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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]

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]

This study presents a new dual-mode propulsion system that utilizes an electrically pumped hydrogen peroxide and kerosene supply system, specifically designed for advanced spacecraft orbit and attitude control. The system incorporates two distinct electric pumps, one for hydrogen peroxide and another for kerosene, each engineered to optimize the propellant feed for the orbit control and attitude control engines. This configuration enables precise propellant management and thrust control, which are critical for the stringent demands of modern space missions. The dual-mode capability of the propulsion system allows for the use of hydrogen peroxide both as a monopropellant in the attitude control engine and as an oxidizer in combination with kerosene in the orbit control engine. This approach simplifies the propulsion architecture while maximizing the efficiency and utility of hydrogen peroxide as a propellant. The electric pumps are essential in providing high-pressure propellant delivery, which is necessary to achieve the desired combustion characteristics and thrust parameters. Performance evaluations of the propulsion system demonstrated significant results. The orbit control engine achieved a combustion chamber pressure of 5.9 MPa and a combustion efficiency of 98.42 %. The attitude control engine, operating at a peak pressure of 4.9 MPa, exhibited a stable pulsative mode, maintaining consistent pressure levels throughout its operational cycle. These performance metrics highlight the effectiveness of the electric pump technology in enhancing propulsion system responsiveness and efficiency. This propulsion system is expected to improve spacecraft maneuverability and reduce operational costs, making it suitable for long-duration space missions. • A versatile dual-mode space propulsion system with electric pump-fed technology. • Advanced electric pumps enable adjustable thrust and precise control. • Confirmed system feasibility and performance through successful ground tests. • Achieved high efficiency and specific impulse in propulsion engines. • Utilized hydrogen peroxide for eco-friendly, non-toxic propulsion. [ABSTRACT FROM AUTHOR]

Ionic liquid electrospray thrusters are notable for their small size, low mass, and high specific impulse, making them ideal for use in micro/nano satellites. This study evaluates the lifetime of an electrowetting ionic liquid electrospray thruster. Experimental results indicate that the decrease in emission current is attributed to electrochemical reactions that increase the curvature of the emitter tips and the increased flow resistance. The online replenishment of insufficient propellant is achieved using an electrowetting device, which restores the emission current. Over the entire test period, the thruster operated for a total of 331.81 h, consuming 2.68083 g of propellant. This study offers valuable insights for improving the lifetime and total impulse of ionic liquid electrospray thrusters. • The lifetime of an electrowetting ionic liquid electrospray thruster is tested. • The thruster has a total operating time of 331.81 h and a mass of propellant consumed is 2.68083 g. • The second online replenishment of emitter propellant is realized by the electrowetting device. • The emission current increases again after propellant replenishment. [ABSTRACT FROM AUTHOR]

Hydroxylamine nitrate (HAN)-based liquid propellants are characterized by high energy content and low toxicity ionic monopropellants, offering great potential application prospects in the space engine. This paper proposes an experimental system for electric ignition and combustion of HAN-based liquid propellant. The energy for decomposition, current, and combustion image of HAN-based liquid propellant are measured under different initial voltage, propellant volume, and the initial pressures in Ar and air atmospheres. The results show that the energy for decomposition decreased with an increase in initial voltage, propellant volume, and initial pressure. An increase in the initial voltage accelerates the reaction speed of HAN-based liquid propellant, and the energy for decomposition is approximately 3 J. An increase in the propellant volume decreases the equivalent resistance in the circuit, increasing the current and enhancing the thermal effect of the propellant. The accumulation of decomposition products and heat under high pressure reduced the energy for decomposition; the energy for decomposition is less than 1.5 J for the pressurization condition. Compared with the Ar atmosphere, the energy for decomposition of HAN-based liquid propellant shows better stability in the air atmosphere. The flow combustion device of HAN-based liquid propellant for the flow state is designed. The HAN-based liquid propellant is successfully ignited for different initial voltages and flow rates. The research results provide a reference for the design of space engine for HAN-based liquid propellant based on the electric ignition method. • Experiments for HAN electric ignition on different initial U , V , P were carried out. • The effects of initial conditions on the ignition energy of HAN was studied. • The comparison of ignition energy in air atmosphere and Ar atmosphere was analyzed. • The propellant flow combustion based-on the electric ignition method was achieved. [ABSTRACT FROM AUTHOR]

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

Transition metal oxides (TMOs) are increasingly viable choices for catalytic applications because of their tendency to exhibit variable oxidation states depending on the chemical environment and the possibility of enhancing surface activity through morphology control. The development of efficient and cost-effective methods for the synthesis of TMOs with meso-to nanoscale morphologies and improved characteristics remains a significant challenge. Herein, we report a template-free synthesis strategy for the preparation of perforated mesoporous manganese(III) oxide microcubes (PMOM) through a morphology-conserved transformation of the precursor method. The formation of PMOM with cubic structures and pore sizes in the range of 20 nm was confirmed using scanning electron microscopy (SEM) and transmission electron microscopy (TEM). The Mn2O3 samples exhibited a Swiss cheese-like porous structure with a particle size of 0.70 to 1.2 μm. Furthermore, the catalytic activity of the synthesized PMOM samples on the thermocatalytic decomposition of ammonium perchlorate (AP), an oxidizer used in solid propellants, was investigated. The transition between Mn3+ to Mn2+ oxidation states facilitates the transfer of electrons, allowing the metal surface to act as an electron acceptor during the oxidation of water molecules. Due to the oxidation–reduction cycle occurring on the surface, a considerable number of electrons are available for promoting surface reactions. Thus, the (211) surface of PMOM enhanced the thermal decomposition of AP, and a plausible mechanism for this catalytic activity was proposed. [ABSTRACT FROM AUTHOR]

New graphene-nickel gallate nanocomplex (G-M-Ni) was designed, synthesized and characterized to deal with the requirement of functional combustion catalyst based on investigations of combustion mechanisms of propellant, structural analysis of catalyst and simulation results of quantum mechanics. The positive effect of G-M-Ni is not only reflected in the significantly increased burning rate, but also reflected in the reduction of friction sensitivity of composite modified double base (CMDB) propellant. The G-M-Ni nanocomplex exhibited significantly enhanced catalytic performances, and the burning rate of propellant at 20 MPa enhanced from 23.15 to 29.07 mm·s−1. Besides, the friction sensitivity of propellant reduced from 64 % to 36 % after addition of 0.5% G-M-Ni. The positive effect of G-M-Ni on CMDB propellant combustion is mainly attributed to its promotion impact on nitroglycerin (NG) pyrolysis, which significantly increased the differential charge density and decreased the decomposition activation energy (from 1.89 eV to 1.38 eV). The gas products produced by the rapid pyrolysis of NG are conducive to the heat feedback between the gas phase and the combustion surface, which makes the propellant have higher burning rate, bright flame, thin dark zone and flocculated quenching surface morphology. The findings in this study confirm the potential of G-M-Ni as a functional combustion catalyst for CMDB propellants, which has guiding significance and inspiration for the development of new combustion catalysts from the view of molecular-level insights. Recently, graphene based functional combustion catalysts have attracted much attention in the field of solid propellants due to their excellent catalytic combustion performance, reduced sensitivity, and improved mechanical properties. Herein, novel graphene-nickel gallate nanocomplex (G-M-Ni) were designed, synthesized and characterized to deal with the to deal with the requirement of functional combustion catalyst based on investigations of combustion mechanisms of CMDB propellant, structural analysis of new material and simulation results of quantum mechanics. By taking graphene as the benchmark, the G-M-Ni nanocomplex exhibited significantly enhanced catalytic performances, and excellent effect on the reduction of friction sensitivity. The relationship between molecular structure of G-M-Ni and combustion characteristics of CMDB propellant were disclosed. The theory-driven design mode of this work may provide new pathways to tune combustion behaviors of CMDB propellants from the view of molecular-level insights. The relationship between molecular structure of G-M-Ni and combustion characteristics of CMDB propellant were disclosed from the view of molecular-level insights. [Display omitted] • Graphene-nickel gallate nanocomplex (G-M-Ni) was first investigated in CMDB propellant. • G-M-Ni is conducive to the combustion and safety performance of CMDB propellant. • Explored the structure-activity relationship of G-M-Ni at the molecular level. • The key catalytic step of catalyst for CMDB propellant combustion was suggested. [ABSTRACT FROM AUTHOR]

A polyol bonding agent (BA) was utilized to enhance the mechanical properties of propellant with low aluminum content. Through scanning electron microscope (SEM) and X‐ray photoelectron spectroscopy (XPS) analysis, it was discovered that BA can physically bond to the surfaces of ammonium perchlorate (AP) and hexogen (RDX). The results of mechanical testing showed that the use of BA can significantly improve the mechanical properties of propellants. Specifically, as the amount of BA increases, the maximum tensile strength (σm) of the propellant also increases, the fracture elongation (εb) initially increases and then decreases, the "dewetting ratio" continuously decreases, and the creep resistance of the propellant also increases. At a BA content of 0.2%, the σm of the propellant reaches 0.6 MPa, the εb exceeds 50%, and the "dewetting ratio" is less than 1.1. [ABSTRACT FROM AUTHOR]

Hollow Cathodes are fundamental components for electric propulsion systems. In recent publications, the performance of a planar C12A7 electride cathode has been presented with promising results. This includes the endurance operation of the cathode, the heaterless ignition cycling, and the general performance mapping over a wide range of parameters. The present publication will add to these publications, presenting results of only one defined set of discharge parameters, allowing a statistical evaluation of repeated discharge operations. Overall, exceptable repeatability of the performance could be evaluated, confirming the overall performance trends observed in previous campaigns. The cathode was operated in self-heating mode using krypton as propellant in a current range of 100 mA to 600 mA. A discharge potential of 30 V with a slight increase for lower discharge currents is reported. Furthermore, means to reduce the variation in the test results by increasing the keeper current have been identified. [ABSTRACT FROM AUTHOR]

Both nanoscience and nanotechnology have undoubtedly contributed significantly to the development of thermite-based nanoenergetic materials (NEMs) with tunable and tailorable combustion performance and their subsequent integration into devices. Specifically, this review article reflects the immense paybacks in designing and fabricating ordered/disordered assembly of energetic materials over multiple length scales (from nano- to milli-scales) in terms of realization of desired reaction rates and sensitivity. Besides presenting a critical review of present advancements made in the synthesis of NEMs, this article touches upon aspects related to various applications concomitantly. The article concludes with the author's summary of the insurmountable challenges and the road ahead toward the deployment of nanoenergetic materials in practical applications. The real challenge lies in the ability to preserve the self-assembly of fuel and oxidizer nanoparticles achieved at the nanoscale while synthesizing macroscale energetic formulations using advanced fabrication techniques both in bulk and thin film forms. Most importantly, these self-assembled NEMs have to exhibit excellent combustion performance at reduced sensitivity to external stimuli such as electrostatic discharge (ESD), friction and impact. [ABSTRACT FROM AUTHOR]

Although various hypergolic propellants like MMH/N2O4 (monomethylhydrazine/dinitrogen tetroxide) are widely used in small rocket engines, there remains a lack of in-depth study conducted on their chemical reactions and spray combustion behaviors. To fill this research gap, a simplified chemical kinetic model that is suitable for three-dimensional simulation was proposed in this paper for MMH/N2O4. Then, numerical investigation was conducted using the Volume of Fluid (VOF) model to explore the transient injection and atomization of MMH/N2O4 impinging jets in a small bipropellant thruster. Also, the instantaneous formation and evolution of the fan-shaped liquid film were analyzed. With the spray distribution determined, the proposed kinetic model and two existing mechanisms were applied to simulate spray combustion and heat transfer within the thruster, respectively, under the Euler–Lagrange framework. According to the research results, the liquid film covered nearly the entire chamber wall with a sawtooth pattern, which protected against the high temperatures of the engine wall. Notably, the two existing mechanisms showed significant errors in predicting temperature changes around the wall due to the excessively simplified reaction pathways. In contrast, the proposed model enabled the accurate prediction of the chamber pressure, wall temperature, and thrust with an error of less than 10%. Given the high accuracy achieved by the proposed numerical method, it provides a valuable reference for the development of advanced space engines. [ABSTRACT FROM AUTHOR]

Background: The lack of visual dynamic spray characterization has made the understanding of the physical processes governing atomization and drug particle formation difficult. This study aimed to investigate the changes in the spray plume morphology and aerodynamic particle size of solution-based pressurized metered-dose inhalers (pMDIs) under different conditions to achieve better drug deposition. Methods: Solution-based pMDIs were studied, and the effects of various factors, such as propellant concentration, orifice diameters, and atomization chamber volume, on drug deposition were examined by analyzing the characteristics of spray plume and aerodynamic particle size. Results: Reducing the actuator orifice and spray area led to a concentrated spray plume and increased duration and speed. Moreover, the aerodynamic particle sizes D50 and D90 decreased, whereas D10 remained relatively unchanged. Decreasing the atomization chamber volume of the actuator led to reduced spray area and an increased duration but a decreased plume velocity. D90 exhibited a decreasing trend, whereas D10 and D50 remained relatively unchanged. Reducing the propellant concentration in the prescription, the spray area and the plume velocity first decreased and then increased. The duration initially increased and then decreased. The values of D50 and D90 showed an initial decreasing followed by an increasing trend, whereas D10 remained relatively unchanged. Conclusions: During the development process, attention should be paid to the changes in the spray area, spray angle, duration, and speed of the spray plume. This study recommended analyzing the characteristics of the spray plume and combining the data of two or more aerodynamic particle size detection methods to verify the deposition in vitro to achieve rapid screening and obtain high lung deposition in vivo. [ABSTRACT FROM AUTHOR]

ABSTRACT Development of new high‐energy materials (HEMs) is one of the thrust areas of research. These compounds are important for various applications such as propellants, gas generations, explosives in mining, construction, civil and military applications, and safety equipment for national security and defense. HEMs based on imidazole frameworks are currently getting research spotlight due to their exceptional detonation performance with optimum sensitivity. This study reports five imidazole derivatives containing azido/nitro/nitrato/fluoro functional groups. These groups are viable options to design promising explosives with moderate sensitivity. Density functional theory (DFT) method is adopted to determine the geometries, thermodynamic properties, detonation properties, and impact sensitivity of the designed molecules. All the compounds have high density in the range of 1.89–2.03 g/cm3 along with high heat of formation. These compounds aid in new strategy to design HEMs with optimum sensitivity. [ABSTRACT FROM AUTHOR]

Nitrocellulose (NC) has been used since the late 1800ʹs in numerous energetic formulations; however, the effect of processing agents (solvents, etc.) on performance has not been thoroughly detailed. For illuminate applications, this is critical as spectral purity and dominant wavelength must meet specifications. In this effort, six processing agents (water, ethanol (EtOH), isopropyl alcohol (IPA), ethyl acetate (EtAc), acetone, and tetrahydrofuran (THF)) are evaluated and the effect they have on the NC fibers and subsequent performance is detailed. XRD and FTIR analysis indicates that the choice of processing agent directly affects the molecular integrity of the NC fibers due to the inhibition of NO2 bonds as well as freeing of hydrogen bonds between polymer chains during solubilization. Time-step FTIR analysis indicated that the weak nitrate ester bond (RCH2-ONO2) is targeted when processed in tetrahydrofuran, ethyl acetate, or acetone, and if heavy solvent interaction occurs, can lead to the autocatalytic decomposition of NC, when heat treated at 110°C. Spectral performance is detrimentally impacted as a result of such interactions. While all six processing agents for red-light illuminant formulations met the military requirements (i.e., spectral purity and wavelength), only ethanol and isopropyl alcohol processed formulations met the military required wavelength of 540 ± 20 nm for green-light illuminants. The IPA processed formulations exhibited the best results with spectral purity values of 69.0% and 94.8% for green and red-light illuminants, respectively. [ABSTRACT FROM AUTHOR]

Mg-based nanofluid fuel is a promising fuel for propulsion system in Mars exploration. In this study, Mg-based nanofluid fuel were prepared by one-step method, and the effect of Mg(magnesium) content, Mg size(68 μm,110 μm,175 μm), and oleic acid (OA) on droplet ignition and combustion under the atmosphere of 90 % CO 2 and 10 % O 2 were investigated. According to the TG-DSC results, 68 m Mg particles under CO 2 had the highest exothermic efficiency of 4472 J/g. The combustion of Mg/OA/kerosene droplets consists of five stages: ignition, d2-law combustion, micro-explosion, vapor flame extinguishment, and agglomerate explosion. The ignition delay time of Mg (68 μm, 5 %)/kerosene is 0.045s, which is half that of kerosene (0.0890s). The agglomerates explode violently and emit a brilliant white light at the end of the Mg-based nanofluid fuel combustion. The explosive effect of Mg agglomerates diminishes as the Mg content increases. The rate of combustion decreases with increasing particle size. In the sample with added oleic acid, Mg (68 μm, 5 %)/OA/kerosene shows the best combustion rate of 0.6320 mm2/s. Finally, a numerical model for the combustion of Mg-based nanofluid fuel droplets is proposed. The predictions are within 2 % of the experimental results. The findings are expected to explore future in-situ utilization of CO 2 on the Mars surface. • Combustion characteristics of a novel Mg-based nanofluid fuel for Mars are studied. • Effects of Mg particle size and content on the combustion properties are analyzed. • The optimal Mg-based nanofluid fuel formulation is Mg (5 wt%, 68um)/OA/kerosene. • A numerical model for the combustion of Mg-based nanofluid fuel droplet is proposed. [ABSTRACT FROM AUTHOR]

Hydrazine's carcinogenic and toxic nature makes its use in space propulsion expensive and necessitates careful launch campaign planning, which may cause delays. One already established alternative is to use energetic salt-based ionic solution propellants. However, the costs remain high because these salts are costly and highly regulated energetic materials. Additionally, propellants based on these energetic salts produce much higher temperatures during decomposition and combustion, as is the case for hydrazine. Consequently, expensive refractory metal-based alloys and high-temperature oxidation-resistant catalytic reactors must be used, which further drives up the costs. Alternatively, hydrogen peroxide may be used, though only in applications with a low specific impulse requirement. This study explores an alternative approach: using nitromethane, a commonly available laboratory solvent. NMP-001 is a nitromethane-based monopropellant that contains additives that allow for low-pressure combustion and reduce sensitivity to mechanical stimuli. Multiple combustion chamber configurations with varying characteristic lengths were tested in rocket combustion experiments. With the 11.5 m L ∗ -configuration, a 15.1 bar low-pressure limit could be determined. The 7.3 and 5.0 m L ∗ -configurations yielded a 16.2 and 28.3 bar low-pressure combustion limits, respectively. These outcomes are significantly below the 30 bar limit reported in the literature. Despite using only a simple heat-sink combustion chamber with a non-sophisticated injector element and a non-optimized propellant formulation, combustion was 83 to 95 % efficient with a C* ranged between 1133 to 1302 m/s. These figures are higher than for hydrogen peroxide and are comparable to the theoretical performance of state-of-the-art green liquid monopropellants. • Existing alternative ionic propellants are costly and highly regulated due to their instability and energetic properties. • Nitromethane offers a cost-effective solution with comparable performance, leading to the development of NMP-001, a nitromethane-based monopropellant with reduced sensitivity to mechanical stimuli. • Tests showed a lower low-pressure combustion limit of 15 ba. This is significantly lower than the limits reported in the literature. • The measured C*-performance for NMP-001 is better than hydrazine and comparable to LMP-103S. [ABSTRACT FROM AUTHOR]

Production of rocket propellants from lunar resources has the potential to significantly reduce the cost of space exploration. Recent research efforts in this area were focused on the extraction of water from icy regolith for conversion into hydrogen and oxygen, a highly efficient rocket bipropellant. However, water is available only in polar regions of the Moon, and its extraction is a challenge. The present paper aims to assess the feasibility of using propellant components that can be obtained from lunar regolith, specifically oxygen, metal alloys, and sulfur. Thermodynamic performance characteristics of rocket engines using these components were calculated over wide ranges of oxidizer-to-fuel mass ratios. It has been shown that the fuel obtained by extraction of oxygen from regolith, i.e., primarily a mixture of metal alloys, exhibits a relatively high specific impulse of up to 250 s. The use of fuel-lean propellants significantly decreases the temperatures, which facilitates cooling and potentially reduces the deposition of condensed products in the engine; at the same time, the expected decrease in the specific impulse is less pronounced. The use of sulfur in rocket engines is less promising from a thermodynamic point of view, but it enables engine designs without a need for feeding metal alloy powders. Among different designs of sulfur-based engines, a hybrid rocket (fuel: metal alloys mixed with sulfur, oxidizer: liquid oxygen) appears to be the most promising. • Rocket propellants can be directly sourced from lunar regolith. • Such regolith-derived fuels can produce a peak specific impulse of up to 250 s. • Sulfur can be used as either an oxidizer or additive to the regolith- derived fuel. • Different configurations for rocket engines using these fuels are presented. [ABSTRACT FROM AUTHOR]

Point cloud data obtained from scanning propellant grains with 3D scanning equipment exhibit positional uncertainty in space, posing significant challenges for calculating the relevant parameters of the propellant grains. Therefore, it is essential to normalize the position of each propellant grain's point cloud. This paper proposes a normalization algorithm for propellant grain point clouds, consisting of two stages, coarse normalization and fine normalization, to achieve high-precision transformations of the point clouds. In the coarse normalization stage, a layer-by-layer feature points detection scheme based on k-dimensional trees (KD-tree) and k-means clustering (k-means) is designed to extract feature points from the propellant grain point cloud. In the fine normalization stage, a rotation angle compensation scheme is proposed to align the fitted symmetry axis of the propellant grain point cloud with the coordinate axes. Finally, comparative experiments with iterative closest point (ICP) and random sample consensus (RANSAC) validate the efficiency of the proposed normalization algorithm. [ABSTRACT FROM AUTHOR]

In this paper, the flow characteristics of a continuous-type liquid–liquid pintle injector are described, focusing on the differential impact of a non-uniform center-flow distribution on single- and bi-injection methodologies as well as the cavitation effect on the spray angle. Using cold-flow experiments, jet-type flows of the center propellant caused by a non-uniform flow distribution were observed during a single injection. This resulted in an augmented pressure drop, as opposed to the flow characteristics of uniform single-film injection. By contrast, bi-injection modalities exhibited a substantial reduction in the pressure drop of the center propellant, underscoring a more equitable flow distribution. Moreover, the occurrence of cavitation in the center propellant was found to markedly affect the spray angle. By considering the injection exit area reduction caused by cavitation, the spray-angle prediction accuracy increased. The findings of this study are expected to reveal the interplay between flow distribution and pressure drop as well as that between cavitation and the spray angle in pintle injectors. Through this understanding, this study provides crucial considerations for the development of more efficient propulsion systems. [ABSTRACT FROM AUTHOR]

In this paper, high cis-1,4 content hydroxyl-terminated polybutadiene (cis-HTPB) with different molecular weights was prepared through the oxidative cracking process using cis-butadiene rubber as raw material. Firstly, this article comprehensively compared the differences between cis-HTPB and conventional I-HTPB in terms of molecular weight distribution, functionality, viscosity, molecular polarity, and other physicochemical properties, which provided effective data support for its subsequent application. In addition, the reaction kinetics study showed that cis-HTPB with isocyanate curing agent has high reactivity, allowing it to be rapidly cured at low temperatures, and the cured elastomers had excellent mechanical properties, with tensile strength and elongation up to 1.89 MPa and 1100%, respectively. It was also found that cis-HTPB has extremely excellent low-temperature resistance, and the glass transition temperature (Tg) of its cured elastomer is as low as -101 ℃. Based on the above studies, cis-HTPB is applied as a binder in composite solid propellants for the first time to investigate its practical performance, and the results indicated that cis-HTPB-based propellants have excellent process and mechanical properties. [ABSTRACT FROM AUTHOR]

This study investigates the combined effects of coated nano copper oxide (NC@nCuO) and polyurethane (PU) modification on the thermal degradation kinetics and combustion properties of ammonium nitrate-based solid composite propellant. Nitrocellulose (NC) is used as a coating for nCuO, forming a core-shell structure, and as a binder in a PU/NC blend. Surface modification of ammonium nitrate was achieved through repeated spray coating. The formulations underwent thorough thermal characterization using thermogravimetry (TG) and differential scanning calorimetry (DSC) analyses. TG analysis showed a significant decrease in thermal degradation temperature due to AN modification with NC@nCuO and the use of PU/NC as a binder. Isoconversional kinetic methods (it-KAS, it-FWO, and VYA/CE) based on DSC tests were used to predict kinetic parameters, revealing a substantial impact of NC-coated nCuO and PU modification on the reactivity of the propellant, reducing the activation energy from 125 kJ/mol (reference) to 99 kJ/mol (nCuO) and 94 kJ/mol (NC@nCuO). Additionally, combustion tests were conducted. The results showed that the introduction of nCuO reduces the ignition delay by 5.25 ms and the combustion duration by 0.06 ms, while NC@nCuO further reduces these by 8.17 ms and 0.12 ms, respectively. Additionally, there is a 23% and 65% increase in flame intensity relative to the reference formulation after the incorporation of nCuO and NC@nCuO, repectively. [ABSTRACT FROM AUTHOR]

Solid rocket motors (SRMs), frequently used as primary boosters in rockets, employ ammonium perchlorate as the oxidizer and aluminum powder as the fuel. Upon combustion, the exhaust products from composite propellants primarily include hydrochloric acid (HCl), carbon monoxide (CO), aluminum oxide, and hydrogen (H2). About 95% of the ammonium perchlorate undergoes conversion to HCl, constituting nearly 20% of the combustion product by weight, significantly contributing to air pollution. To address this environmental concern, we have developed a reusable MoS2 aerogel specifically engineered for absorbing HCl gas. The fabrication process is straightforward and cost‐effective, utilizing nontoxic reagents and standard equipment, enhancing its practicality. Characterization of the aerogel was conducted using Scanning Electron Microscopy (SEM), Transmission Electron Microscopy (TEM), X‐ray Diffraction (XRD), X‐ray Photoelectron Spectroscopy (XPS), and Brunauer–Emmett–Teller (BET) analysis. The aerogel demonstrated a high absorptive capacity, achieving 65.88% absorption efficiency for HCl gas, which is noteworthy. Moreover, the aerogel samples exhibited exceptional recyclability over five cycles, highlighting their suitability for applications requiring HCl gas absorption. This research lays the foundation for advancing and scaling up innovative nanomaterial‐based aerogels, offering versatile solutions for a wide range of contemporary applications. [ABSTRACT FROM AUTHOR]

To improve the reinforcement effect between a binder and high solid filler in a propellant formula, grafting the bonding group into the binder to form a neutral polymeric is a practically novel approach to improving the interface properties of the propellant. In this work, a glycidyl azide polyol energetic thermoplastic elastomer binder with a –CN bonding group (GAP–ETPE) was synthesized, and the mechanical and thermal decomposition mechanism of GAP–ETPE with Hexogeon (RDX) model propellants were studied. The stress–strain results indicated that the tensile strength and strain of GAP–ETPE/RDX model propellants were 6.43 MPa and 32.1%, respectively. DMA data showed that the storage modulus (E') of the GAP–ETPE/RDX model propellants could increase the glass transition temperature (Tg) values, those were shifted to higher temperature with the increase in filler RDX percentages. TG/DTG showed the four decomposition stages of the decomposition process of the GAP–ETPE/RDX model propellants, and the thermal decomposition equation was constructed. These efforts provide a novel method to improve GAP–ETPE/RDX propellants mechanical property, and the thermal decomposition behavior of GAP–ETPE/RDX propellants also provided technical support for the study of propellant combustion characteristics. [ABSTRACT FROM AUTHOR]

Gradiently denitrated gun propellant (GDGP) prepared from nitrocellulose-based propellant by denitration strategy has the advantages of high combustion progressivity and storage stability, but still presents the problems of difficult ignition and lower oxygen balance than that of the raw propellant. In addition, the low utilization efficiency of uniformly added ablation inhibitors in conventional propellant formulations can also lead to harmful shooting phenomena. In this paper, nano-zinc oxide (ZnO) with high thermal conductivity as one of the ablation inhibitors was loaded on the surface of GDGP for the first time. FTIR, Raman, XRD, SEM, and XPS characterizations confirmed the preparation and structural integrity of the GDGP and gradiently denitrated gun propellant-ZnO composite (GDGP@ZnO) with ZnO in nano-size and adjustable loading concentration. The test results showed that GDGP@ZnO exhibited higher combustion progressivity than raw propellant, and the appropriate loading concentration of ZnO could enhance the heat conduction across the surface of the gun propellant and improve the ignition performance. The ablation inhibitor loaded on the surface of the gun propellant can provide an ablation inhibition effect while reducing the dosage. This work provides a new idea for the design and fabrication of a novel multifunctional integrated gun propellant with high combustion progressivity, easy ignition and low ablation. [ABSTRACT FROM AUTHOR]

Using the possibilities of modern science and technology, it can be said that the presented mathematical model has been formulated without significant simplifications, and its (numerical) solution itself will be performed with high accuracy. The main purpose of the work was to create a tool supporting the calculation of the main problem of internal ballistics for barrel propellant systems in order to achieve digital solutions as close as possible to the results obtained using experimental ballistics methods. The scientific hypothesis of the work postulates that the physical model of internal ballistics formulated in Lagrange coordinates will allow to solve the main problem of internal ballistics of barrel systems in a digital way, obtaining such solution results that will satisfactorily reflect the solution results obtained using experimental ballistics methods. The results of digital simulations of physical phenomena of solutions to the main problem of internal ballistics of barrel weapons were compared with the results of experimental tests and the degree of agreement was determined. [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]

This study conducted a detailed examination of the characteristics of hydroxyl‑terminated polybutadiene (HTPB) doped with potato starch nitrate (NPS) as a high-energy-density binder. The investigation specifically focused on evaluating the impact of NPS doping on the chemical structure, energetic attributes, and thermokinetic behavior of HTPB-NC binder systems. The findings demonstrated advantageous features, including favorable chemical compatibility, improved material density accompanied by remarkable thermal stability, substantial energy release, and high chemical reactivity. These findings contribute to the understanding of HTPB-NPS composite and offer evidence for its remarkable features, indicating promising prospects for its utilization in the development of advanced formulations for high-energy-density propellants in future applications.

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]

Carbon nanotubes (CNTs) loaded with copper oxide (CuO) as the combustion catalysts for solid propellants are able to increase the combustion rate and reduce the pressure index of propellants. The dispersibility of CNTs in water is one of the most important factors influencing the loading of CuO. In this paper, 1-dodecyl−3-methylimidazolium chloride ([C12C1Im]Cl) was prepared and used as the pretreatment agent of CNTs. The Nuclear magnetic resonance (NMR) results indicated the successful synthesis of [C12C1Im]Cl, which contains a positively charged imidazolium hydrophilic headgroup, enabling it as a cationic surfactant for the surface pretreatment of CNTs. The properties of the [C12C1Im]Cl-pretreated CNTs were compared with those pretreated with sodium dodecyl sulfate (SDS) and mixed acids (with a volume ratio of concentrated sulfuric acid to concentrated nitric acid = 3:1). X-ray diffraction (XRD) results showed that the crystalline structures of the CNTs were unchanged by using all of the three pretreatment methods. Scanning electron microscope (SEM) morphology analysis showed that the CNTs pretreated with [C12C1Im]Cl had the clearest tubular shape, the smoothest walls and the best dispersion in water without agglomeration. The impact of the pretreatment methods on the loading of CuO onto the surface of CNTs was also illustrated. It was concluded that by pretreatment with [C12C1Im]Cl, it led to the best CuO dispersion on the CNTs surface, the smallest CuO grain size and the largest CuO content. The study on the loading mechanism indicated that CuO could be successfully loaded on CNTs only when the pH > 4.6014, given the temperature is higher than 75°C. [ABSTRACT FROM AUTHOR]

A circular waveguide electron cyclotron resonance plasma thruster prototype driven by microwaves at 5.8 GHz (80–300 W) is characterized. The magnetic field is generated by a combination of Sm-CoYXG32 magnets and an electromagnet, which enables the tuning of the resonance position and magnetic nozzle shape. The main plasma plume properties are analyzed by using electrostatic probes when the mass flow rate (Xenon), microwave power, electromagnet current, and propellant injector design are varied. An estimation of the propulsive performance of the device is also presented. Results show that a single radial injector hole is not sufficient for a symmetric ion current profile and that magnetic nozzle shape and strength tuning can significantly affect the divergence angle and thruster floating potential. A utilization efficiency of up to 70% and electron temperatures of up to 16 eV have been measured. [ABSTRACT FROM AUTHOR]

Through the perspective of differential equations, the report "Rocket Science Unveiled" explores the amazing invention of rocket propulsion. In order to study, comprehend, and forecast the behavior of rocket engines, differential equations are essential. In order to better understand and analyze this intricate anomaly, the report aims to investigate the underlying mathematics of rocket propulsion and how differential equations work. We apply the differential equation to clarify the fuel consumption and thrust generation rates. In addition, we utilize Newton's rule of motion to explain the relationship among thrust, mass, and acceleration. Working on this study allowed us to discover the anticipated outcome for both position location and spacecraft position determination. For iterative operations, we used Euler's approach because the analytical calculation of differential equations is complicated, we used Euler's method for iterative operations. Knowing the rocket's initial or previous value allows us to locate or establish its placements with ease.

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]

The application of lightweight carbon materials and their composites in solid propellants, especially graphene and its derivatives as combustion catalysts in the combustion performance regulation of propellants have attracted attention in the field of aerospace propulsion. The energetic composite µAl/NGO was prepared using nitrated graphene oxide (NGO) and micron-aluminum powder (µAl) as raw materials by solution method. The structure and morphology of the composite was characterized by elemental analysis, Fourier-transform infrared spectroscopy, X-ray diffraction, X-ray photoelectron spectroscopy and scanning electron microscopy. The thermal properties of the composite were analyzed by TG-DSC test results. The results show that µAl is uniformly distributed on the surface of NGO, and the crystal structure of µAl is not changed. In the temperature range of 300–450°C, µAl/NGO explodes, and it is a rapid exothermic reaction with a peak temperature of 364°C. In this temperature range, the decomposition of NGO and the oxidation of µAl occurred simultaneously by TG results. Compared with the enthalpy of oxidation exothermic process of the pure µAl at 557°C (2274 J/g), the enthalpy of decomposition and oxidation exothermic processes of µAl/NGO is increased to 10484 J/g, which is nearly 5 times higher than that of the pure µAl. The result of μAl/NGO catalytic behavior on the thermal decomposition of ammonium perchlorate (AP) indicate that μAl/NGO has little effect on the low temperature thermal decomposition of AP, but the peak of high temperature thermal decomposition advances from 404 to 364°C. The kinetic analysis results show that the activation energy of the high-temperature exothermic reaction of AP is increased by nearly 100 kJ/mol. [ABSTRACT FROM AUTHOR]