Search

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]

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]

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]

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]

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.

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 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]

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]

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 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]

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]

In the contemporary landscape characterized by heightened competitive pressures, adept maintenance management assumes a pivotal and strategic role in cost mitigation and profit augmentation. This paper applies the risk-based maintenance (RBM) methodology to the hydraulic system of a remotely operated vehicle (ROV) owned by an oil and gas services company. To achieve this, it was necessary to gather information on the system failures, followed by determining the critical subsystems. By analyzing the frequency and severity of identified failures, a risk matrix could be constructed. Using the "5 Whys" technique, an analytical framework was used to uncover the root causes of the failures observed in the handlers and propellant subsystems. Subsequently, a risk matrix is formulated again after the implementation of the RBM and the mitigation measures adopted by the authors to elucidate the discernible impact of the maintenance methodology on improving the availability of the ROV and the incidence of failures outlined in the study. [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]

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]

Hypergolic propellants rely on fuel and oxidizer that spontaneously ignite upon contact, which fulfill a wide variety of mission roles in launch vehicles and spacecraft. Energy‐rich carboranes are promising hypergolic fuels, but triggering their energy release is quite difficult because of their ultrastable aromatic cage structure. To steer the development of carborane‐based high‐performance hypergolic material, carboranylthiolated compounds integrated with atomically precise copper clusters are presented, yielding two distinct isomers, Cu14B‐S and Cu14C‐S, both possessing similar ligands and core structures. With the migration of thiolate groups from carbon atoms to boron atoms, the ignition delay (ID) time shortened from 6870 to 3 ms when contacted with environmentally benign oxidizer high‐test peroxide (HTP, with a H2O2 concentration of 90%). The extraordinarily short ignition ID time of Cu14B‐S is ranking among the best of HTP‐active hypergolic materials. The experimental and theoretical findings reveal that benefitting from the migration of thiolate groups, Cu14B‐S, characterized by an electron‐rich metal kernel, displays enhanced reducibility and superior charge transfer efficiency. This results in exceptional activation rates with HTP, consequently inducing carborane combustion and the simultaneous release of energy. This fundamental investigation shed light on the development of advanced green hypergolic propulsion systems. [ABSTRACT FROM AUTHOR]

Hydrofluorocarbons (HFCs) are powerful anthropogenic greenhouse gases (GHGs) with high global-warming potentials (GWPs). They have been widely used as refrigerants, insulation foam-blowing agents, aerosol propellants, and fire suppression agents. Since the mid-1990s, emissions of HFCs have been increasing rapidly as they are used in many applications to replace ozone-depleting chlorofluorocarbons (CFCs) and hydrochlorofluorocarbons (HCFCs) whose consumption and production have been phased out under the Montreal Protocol (MP). Due to the high GWP of HFCs, the Kigali Amendment to the MP requires the phasedown of production and consumption of HFCs to gradually achieve an 80 %–85 % reduction by 2047, starting in 2019 for non-Article 5 (developed) countries with a 10 % reduction against each defined baseline and later schedules for Article 5 (developing) countries. In this study, we have examined long-term high-precision measurements of atmospheric abundances of five major HFCs (HFC-134a, HFC-143a, HFC-32, HFC-125, and HFC-152a) at Gosan station, Jeju Island, South Korea, from 2008 to 2020. Background abundances of HFCs gradually increased, and the inflow of polluted air masses with elevated abundances from surrounding source regions were detected over the entire period. From these pollution events, we inferred regional and country-specific HFC emission estimates using two independent Lagrangian particle dispersion models and Bayesian inversion frameworks (FLEXPART-FLEXINVERT + and NAME-InTEM). The spatial distribution of the derived "top-down" (measurement based) emissions for all HFCs shows large fluxes from megacities and industrial areas in the region. Our most important finding is that HFC emissions in eastern China and Japan have sharply increased from 2016 to 2018. The contribution of East Asian HFC emissions to the global total increased from 9 % (2008–2014) to 13 % (2016–2020). In particular, HFC emissions in Japan (Annex I country) rose rapidly from 2016 onward, with accumulated total inferred HFC emissions being ∼ 114 Gg yr -1 , which is ∼ 76 Gg yr -1 higher for 2016–2020 than the "bottom-up" (i.e., based on activity data and emission factors) emissions of ∼ 38 Gg yr -1 reported to the United Nations Framework Convention on Climate Change (UNFCCC). This is likely related to the increase in domestic demand in Japan for refrigerants and air-conditioning-system-related products and incomplete accounting. A downward trend of HFC emissions that started in 2019 reflects the effectiveness of the F-gas policy in Japan. Eastern China and South Korea, though not obligated to report to the UNFCCC, voluntarily reported emissions, which also show differences between top-down and bottom-up emission estimates, demonstrating the need for atmospheric measurements, comprehensive data analysis, and accurate reporting for precise emission management. Further, the proportional contribution of each country's CO 2 -equivalent HFC emissions has changed over time, with HFC-134a decreasing and HFC-125 increasing. This demonstrates the transition in the predominant HFC substances contributing to global warming in each country. [ABSTRACT FROM AUTHOR]

In this study, an experimental apparatus was designed to measure the thermochemical erosion properties of gun steel samples. Three tests were conducted in propellant gas to analyze the surface, cross-sectional morphology, and major element distribution. The results show that the specimen surface undergoes significant erosion by the hot gas, forming a porous white layer with a low melting point. The white layer's thickness and degree of damage increase with prolonged erosion time. During erosion, abundant amounts of carbon and oxygen migrate to the surface and bonded with iron. Carbonation reduces impact toughness, while oxidation increases metal volume, resulting in a more porous white layer. Consequently, many microcracks form within the white layer, which grow and intersect with deepening erosion. These findings suggest that surface erosion can lead to fracturing and breakage of the material, ultimately increasing exposure of the substrate to gas and accelerating the erosion rate. [ABSTRACT FROM AUTHOR]

This work represents a first attempt to include the complex variety of electron- molecule processes in a full kinetic particle-in-cell/test particle Monte Carlo model for the plasma and neutral gas phase in a Hall thruster. Particular emphasis has been placed on Earth's atmosphere species for the air- breathing concept. The coupling between the plasma and the gas phase is self-consistently captured by assuming the cold gas approximation and considering gas-wall and gas recycling from the walls due to ion neutralization. The results showed that, with air molecular propellants, all the most relevant thruster performance figures degraded relative to the nominal case using Xe propellant. The main reasons can be ascribed to a reduced ionization cross-section, a larger gas ionization mean free path due to lighter mass air species, and additional electron collisional power losses. While vibrational excitations power losses are negligible, dissociation and electronic excitations compete with the ionization channel. In addition, for molecular oxygen, the large dissociation leads to even faster atoms, further reducing their transit time inside the discharge channel. Future studies are needed to investigate the role of non-equilibrium vibrational kinetics and metastable states for stepwise ionization. [ABSTRACT FROM AUTHOR]

This study presents the first experimental demonstration of a Hall Effect Thruster operating with water electrolysis-derived constituents: oxygen for the anode (thruster) and hydrogen for the cathode, both from pressurised bottles. In contrast to prior research where the thruster was operated with oxygen to the anode and xenon/krypton to the cathode, this investigation uses an advanced set-up that safely allowed the combined use of pressurised oxygen and hydrogen. Firstly, a comparative assessment between the established plasma filament cathode (utilised in previous research) and a novel L a B 6 cathode (referred to as Hydrocat), both krypton-fuelled, is conducted. Results show a different IV curve for these cathodes, but negligible disparities in thruster performance for fixed discharge power and anode mass flow. Thrust, anode specific impulse, thrust-to-power ratio, and both, anode and total efficiency of the system using oxygen (anode) and hydrogen (cathode) are then presented and compared against oxygen (anode) and krypton (cathode). The cathode successfully operated on hydrogen, with a reduction in thrust performance of under 10% compared to when krypton is used as the cathode propellant. This performance difference can be attributed to different factors, such as the reduced efficiency of the cathode running with hydrogen due to a higher ionisation potential and the cathode propellant's contribution to thrust, more significant when using heavier propellants such as xenon/krypton compared to hydrogen. The impact of changing the hydrogen-to-oxygen ratio is also assessed and shown to have no impact on the thruster's performance. The most optimal performance was obtained at the highest discharge power tested ( P d = 1953 ± 1 W), using a hydrogen:oxygen ratio of Φ = 1:25. This operating point corresponded to a thrust of 20. 9 ± 0. 2 mN, anode specific impulse of 2130 ± 20 s, thrust-to-power ratio of 10. 7 ± 0. 5 mN/kW, anode thrust efficiency of 11. 2 ± 0. 3 %, and total efficiency of 10. 1 ± 0. 3 %. • First demonstration of a Water Electrolysis fuelled Hall Effect Thruster. • Performance comparison between a krypton-fuelled and a hydrogen-fuelled cathode. • Thrust, specific impulse, anode efficiency and total efficiency metrics. • Influence of the water electrolysis stoichiometric ratio on the performance. • Oxidation studies when using oxygen and hydrogen plasmas. [ABSTRACT FROM AUTHOR]

Slow-burn propellants are crucial for aerospace advancement. This study focuses on the development of slow-burn propellants by replacing a small amount of AP (3–9 %) with various amine-based burning rate suppressants (BRSs). The consistency between thermal analysis and combustion test results indicates a direct link between the main exothermic decomposition process and combustion performance, with changes in onset temperature positively correlated with decreases in combustion parameters. Upon adding 3 % oxamide, the onset temperature of the main decomposition peak in DSC shifted backward by 16.2 °C, significantly higher than the formulations containing ammonium sulfate at 2.6 °C and ammonium oxalate at −0.8 °C. As the content of oxamide further increased from 3 to 9 %, the ignition delay time increased simultaneously by 0.300–1.638 s, atmospheric burning rate decreased by 14.01–45.45 %. These values represent the maximum among formulations with the same content of the BRSs, mainly attributed to its higher heat absorption capacity. • The onset temperature shift of the high-temperature decomposition peak, as the primary energy release phase, is positively correlated with the extent of combustion performance decline. • The introduction of three types of amine-based burning rate suppressants does not affect the light-absorbing properties of the propellants, but significantly influences the ignition behavior. • Infrared cameras have verified that certain particles of ammonium sulfate exit the burning surface alongside the gas before complete decomposition, resulting in a diminished effect on the reduction of the burning rate. [ABSTRACT FROM AUTHOR]

The present study examines the advancements of a promising low-toxicity hypergolic propellant combination called HIP_11 for in-space applications, developed as alternative to common hypergolic toxic propellants. The fuel is based on an ionic liquid, storable, stable and simple to handle at ambient conditions, developed at the Institute of Space Propulsion, German Aerospace Centre (DLR). The compound shows hypergolicity behaviour when in contact with Hydrogen Peroxide and is a promising substitution to conventional hypergolic propellants. The present work describes the advancements in the development of HIP_11 through a dedicated experimental investigation of various components, analysing and investigating the performances. The experiments are based on a small modulable thruster that allowed to study the efficiency and stability of combustion of the propellant combination while varying various design parameters of the thruster. Specifically, the effects of different injector designs, as well as variations in combustion chamber shape, characteristic length, and operating pressure have been thoroughly examined and analysed. • Experimental testing of a novel green hypergolic propellant for in-space propulsion. • Investigation of multiple design alternatives for critical components. • Analysis of multiple injector and combustion chamber designs in hot-firing tests. • Evaluation of combustion performance in terms of combustion efficiency and stability. • Enhanced combustion stability noted for short and slender combustion chambers with impinging injectors. [ABSTRACT FROM AUTHOR]

Pintle injectors have attracted much attention in recent years for their excellent throttling and combustion performance. However, the combustion characteristics of cryogenic propellant on pintle injectors with varying operating conditions remain unclear. A LOX/kerosene throttleable rocket engine with a fuel-centered liquid-liquid pintle injector was designed and its pressure drop and combustion characteristics were tested. The pintle engine achieved stable combustion with the total flow rate and oxygen-fuel mixture ratio varying from 59 to 559 g/s and from 0.73 to 1.02. A calculation model for the gas ratio in the LOX channel of the pintle injector was established. The deviation of measurements from the calculated values did not exceed ±10%. The vaporization of LOX from the pintle injector in the hot test was calculated and the positive correlation between gas ratio and combustion efficiency was analyzed. For the phenomenon of low-frequency oscillation of pressure in the hot tests, the source and development process of the oscillation were studied by FFT. The primary frequency and amplitude of low-frequency oscillation had been discovered to exhibit a positive correlation with the gas ratio. Low pressure drop of injection was the main cause of low-frequency oscillation. The strong vaporization of LOX under low operating conditions made the combustion efficiency remain in a high range even if low-frequency oscillation was observed. This work can provide a reference for the design of pintle injectors for cryogenic liquid rocket engines that can be deeply throttled. • A model for calculating the amount of LOX vaporization in the pintle injector was presented. • The relationship between combustion efficiency and gas ratio in the LOX channel was analyzed by partial correlation. • The source and development process of the low-frequency oscillation were studied by FFT. [ABSTRACT FROM AUTHOR]

The formulation fundamentally determines the combustion characteristics and agglomeration behavior of propellants. This study compared the effects of four factors on the combustion, agglomeration, and combustion efficiency of nitramine solid propellants: the aluminum size, the ammonium perchlorate size, the nitramine size, and the mass ratio of ammonium perchlorate to nitramine. The comparison was conducted through a series of experimental methods including burning rate tests, collection of the condensed-phase combustion products, chemical titration, and physicochemical analysis. In terms of burning rate, reducing the particle sizes of aluminum powder and ammonium perchlorate, increasing the nitramine size, and increasing the mass ratio of ammonium perchlorate to nitramine all effectively enhance the burning rate. The influence of these four factors on burning rate, in descending order of impact, is: the mass ratio of ammonium perchlorate to nitramine > the ammonium perchlorate size > the aluminum size > the nitramine size. Regarding agglomeration and combustion efficiency, increasing the particle sizes of aluminum and nitramine, increasing the mass ratio of ammonium perchlorate to nitramine, and decreasing the ammonium perchlorate size can effectively inhibit agglomeration and enhance combustion efficiency. Among these, the ammonium perchlorate size has the greatest effect on enhancing the combustion efficiency, potentially increasing it by 16%. The nitramine size has the weakest impact on combustion efficiency, with a potential increase of 5.5%. The experimental results of this study can provide references for a deeper understanding of the combustion mechanisms of nitramine solid propellants, the establishment of combustion models, and the realization of formula optimization. • Formulation effect on combustion and agglomeration of propellants are studied. • The mass ratio of AP to nitramine has the greatest effect on the burning rate. • The AP size has the greatest effect on enhancing the combustion efficiency. • The comprehensive mechanism of formulation on propellants is proposed. [ABSTRACT FROM AUTHOR]

The application of hexanitrohexaazaisowurtzitane (HNIW) as an oxidizer in solid propellants aligns with the pursuit of high-energy materials. However, the phase transformation behavior and high impact sensitivity of HNIW are its limitations. Due to the strong adhesion and mild synthesis conditions, polydopamine (PDA) has been employed to modify HNIW. However, the method suffers from a slow coating process and a non-ideal coating effect under short reaction time. Herein, oxygen-accelerated dopamine in situ polymerization coating method was developed. It was found that oxygen not only reduced the coating time but also contributed to forming a dense and uniform PDA layer. HNIW@PDA coated in oxygen for 6 h exhibited the most favorable performance, with a delay of 20.8 °C in the phase transition temperature and a reduction of 145.45% in the impact sensitivity. The -OH groups on the surface of PDA enhanced the interaction between HNIW and polymer binders, resulting in a 20.36% reduction in the dewetting percentage. The lower content of PDA in HNIW@PDA (1.17%) resulted in minimal variation in the heat of explosion for HNIW@PDA-based HTPB propellant (6287 kJ/kg) in comparison to HNIW-based HTPB propellant (6297 kJ/kg). Hence, HNIW@PDA-based propellants are expected to offer an alternative with promising safety and mechanical performance compared to existing HNIW-based propellants, thus facilitating the application of HNIW in high-energy propellants. This work presents a low-cost method for efficiently inhibiting the phase transformation of polycrystalline explosives and reducing the impact sensitivity. It also offers a potential approach to enhance the interfacial interaction between nitro-containing explosives and polymer binders. [ABSTRACT FROM AUTHOR]

Context: Explosive properties of novel potential high energy density materials of a series of 1,2-diazete-based molecules with trinitromethyl functional group were investigated computationally. All the sixty seven molecules were optimised to obtain their molecular geometries and electronic structures. Electrostatic potential analysis was also carried out in the determination of different parameters. The calculations indicate that the majority of the compounds have high positive heat of formations, high density and good detonation performance greater than that of traditional energetic materials like RDX and HMX. They are also having comparable values of impact sensitivity. These features promise their potential to be used as energetic materials for future applications. Most of the designed molecules are having high positive oxygen balance values so that the study of these molecules can also be extended as potential candidates for oxidisers in solid propellants. Methods: Optimisation and vibrational frequency analysis of the studied molecules were done with density functional theory using B3LYP/aug-cc-pVDZ as the basis set to zero imaginary frequencies using Gaussian 09. Electrostatic potential analysis were carried using the Multiwfn program. [ABSTRACT FROM AUTHOR]

Coating modification is an important way to enhance the reactivity of aluminum powder. In this paper, ammonium perchlorate and aluminum powder were assembled into energetic microunits by liquid deposition method. Spherical particles with AP as shell and ultrafine aluminum powder as the core (Al@AP) were gained. The micromorphology results show that the coated particles are about 5 mm, and the coating layer is evenly distributed on the outer surface of aluminum powder, indicating a complete coating. The energetic microunits were implanted into the nitrate ester plasticizing adhesive system (NEPE) as solid phase fillers. The effect of filler on the rheological properties, safety, mechanical properties, thermal reaction and energy properties of the system was analyzed by comparing with the raw aluminum filler. The test results show that the rheological properties, mechanical properties and pressure index of NEPE containing system Al@AP meets the requirements of solid propellant charging. Compared with Al based propellant, the mechanical sensitivity and thermal sensitivity are decreased, the safety is better, and the explosion heat of the propellant is increased by 7:8%. The engine test shows that the specific impulse is increased by 1.2 s. Al@AP can improve the energy output and safety of NEPE propellant, and has potential application prospects in high-energy propellants. [ABSTRACT FROM AUTHOR]

To achieve the controllable release of energy of nitrocellulose-based propellants, this paper combines the cellulose-based nanocomposites aqueous coating (Surelease®-NC) with fluidized bed coating equipment to successfully prepare the coated spherical propellant for the first time. The effects of fluidized bed coating temperature, air velocity, flow speed and atomization pressure on the adhesion rate, coating integrity and coating uniformity of the coated spherical propellant were investigated, and the preparation of coated spherical propellant with homogeneous size and structural integrity was achieved for the first time. The microscopic morphology, chemical structure, water vapor adsorption behavior, combustion performance, and ageing resistance property of the coated spherical propellant were systematically investigated by, Fourier transforms infrared spectroscopy (FTIR), Micro confocal raman spectrometer, field scanning electron microscopy (SEM), dynamic vapor adsorption techniques, and closed bomb test, confirming the surface core-shell structure and the tightly bonded interfacial structure of coated spherical propellant. Meanwhile, the coated spherical propellant has good hygroscopic, excellent progressive burning and long storage stability. [ABSTRACT FROM AUTHOR]

Due to technical limitations and the characteristics of the launch power system propellant, its support capability is affected by many factors and links. In the process of mission support, it is necessary to consider both its quality characteristics and the safety during the support process, which bring difficulties to the accuracy of its support capability evaluation. In this regard, the evaluation indexes system is constructed according to the requirements of total quality management theory, process safety management theory and equipment support characteristics. In the process of setting the indexes system, the technique for order preference by similarity to an ideal solution method (TOPSIS method) is used to screen the indexes to make indexes setting more scientific. The subjective weights obtained by the order relation analysis method (G1 method) and the objective weights obtained by the entropy weight method (EW method) are combined by the game theory combination weighting method to obtain the indexes combination weights, which reduce the error of weight calculation. The two-dimensional cloud model is used to evaluate support capability of the launch power system propellant from the two dimensions of quality management capability and safety management capability, which can reflect the support capability more comprehensively and accurately. Through example verification, the closeness of support capability of the launch power system propellant to the evaluation grades of 'excellent', 'good', 'average', 'poor' and 'very poor' is calculated to be 0.2333, 5.1169, 0.3579, 0.1821 and 0.1041, respectively, indicating that the evaluation results are closer to good. Finally, compared with the existing one-dimensional cloud model evaluation method, the superiority of this research method is highlighted, which has a certain guiding role in improving its support capability. [ABSTRACT FROM AUTHOR]

As a new type of green propellant, ammonium dinitramide (ADN)-based energetic propellants have wide application value and development potential in the field of space propulsion. This paper delves into the intricate impact of varying initial temperatures, pressures, and propellant component ratios on critical parameters, including temperature, combustion rate, and heat release, in the straight nozzle of an ADN-based propellant. The findings indicate that an elevation in both initial temperature and ADN ratio expedites the thermal decomposition rate of ADN, thereby elevating the average temperature in the nozzle. However, the elevation in initial temperature has a negative effect on the overall rise amplitude of average temperature. Furthermore, the initial pressure setting is crucial in determining whether the oxidation reaction of the fuel CH3OH occurs in ADN propellants. When the initial pressure is greater than 10 atm, CH3OH is completely consumed, and the final average temperature is about 2650 K, which increases by 558.89% compared with that at 1 atm. Our work aims to provide theoretical guidance and practical optimization strategies for enhancing propellant performance and optimizing thruster structure design. [ABSTRACT FROM AUTHOR]