Search Results (13)

This review provides a state-of-the-art and up-to-date analysis of the design and development of green monopropellant thrusters based on hydrogen peroxide (H₂O₂) and hydroxyl ammonium nitrate (HAN) as high-energy compounds for reaction control maneuvering of satellites. In summary, we introduce the new concept of Green Liquid Propellants (GLPs) that can serve as eco-friendly alternatives to conventional hydrazine thrusters. GLPs offer several advantages, including low toxicity, acceptable thermal decomposition and combustion behaviors, low onset temperatures of decomposition, stability, and long-term storability, compared to hydrazine. H₂O₂ exhibits a low onset temperature; however, its storability does not match that of hydrazine. On the other hand, HAN boasts excellent storability; however, it comes with a higher onset temperature when compared to hydrazine. This review provides critical insights into the recent advancements in H₂O₂ and HAN thrusters, along with an examination of the corresponding catalysts. The focus is on their application for the long-term maneuvering of satellites. We have chosen H₂O₂ and HAN formulations to focus on these two GLPs due to their extensive use by various space agencies worldwide. Moreover, the future directives of both selected green propellants were discussed for potential applications. Finally, the choice between H₂O₂ and HAN depends on the specific requirements of the propulsion system, taking into account factors such as performance, environmental impact, safety, and operational considerations. Each propellant has its advantages and challenges, and ongoing research aims to address some of the limitations associated with these green propellants. Full article

Hydrogen peroxide (High Test Peroxide, HTP) emerges as a promising candidate for green space propulsion applications due to its lower toxicity compared to liquid conventional propellants such as hydrazine and nitrogen tetroxide. This study aims to optimize the performance and reliability of HTP monopropellant thrusters, focusing on catalyst bed stability, efficiency, and durability during extended steady-state operations. Key parameters, including catalyst bed packing, pellet size, bed load, and HTP concentration, were investigated in this study for their impact on the steady-state performance, using the pressure loss across the catalyst bed as an indicator of catalyst deterioration. Results indicate that an optimal pressure drop of 1–1.5 bar across the catalyst bed provides optimal stability and durability. To evaluate transient characteristics, effects of bed load, HTP concentration, and pre-heating temperature on thruster response times were investigated. Following the optimization process, a lifetime test with an HTP throughput of 6 kg was conducted to monitor performance variations over time. Additionally, the blowdown characteristics of the thruster were analyzed to assess performance under end-of-life conditions. The experiments in this study demonstrate that HTP monopropellant thrusters are viable candidates for reliable space missions, particularly for long-duration operations such as station-keeping maneuvers. Full article

In a worldwide scenario which sees an increasing number of small satellite launches, novel mission concepts may be unlocked providing the spacecrafts with the very precise and rapid maneuvering capability that electric thrusters cannot guarantee. In this context, chemical thrusters appear to be a possible solution. This work aimed to experimentally study and solve the problem of ignition for 10 N hybrid rockets based on hydrogen peroxide. Firstly, the study analyzed the performance of a monopropellant engine capable of functioning as a hybrid injection system. In particular, the effects of the liquid mass injected, the initial temperature, and the supply pressure on the pulsed engine performance were experimentally investigated. The injected mass showed a greater impact on the performance with respect to the starting chamber temperature and injection pressure. This thruster also showed a good potential for space applications. In the second part of the work, the objective was to find an ignition procedure that reduced propellant consumption and eliminated the need for a glow plug. This is important because the electrical power consumption in real applications significantly affects other subsystems and is undesirable for chemical engines. Different ignition procedures were tested to emphasize their respective advantages and disadvantages, and the findings indicated that the concept of pulsed preheating is feasible with only a small amount of propellant consumption, while substantially decreasing the ignition duration from approximately 45 min to a maximum of just 3 min. Finally, similar ignition procedures were adopted using different fuels. The results showed that PVC and ABS, under the same operating conditions, ignite more easily than HDPE, which requires an oxidizer consumption approximately double that of the other two fuels. Considerations about the effect of chamber pressure and oxidizer mass flow rate on engine ignition were also included. Full article

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. Full article

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. Full article

The development of propulsion systems based on green propellants, as an alternative to hydrazines, has been gaining interest within the space community. The study of Ammonium Dinitramide (ADN)-based liquid monopropellant, which is low-toxic and can deliver high performance, is the focal point of interest for Space Solutions Co., Ltd., Daejeon, Republic of Korea. A 1N ADN-based propulsion system was designed to evaluate the performance of the propellant. By combining a thermal and catalytic bed in a reactor, the performance of the propellant was examined in a designed thruster (chamber pressure of 11 bar). A total of 16 tests, with pulse mode experiments, were conducted; the accumulated firing time was 285 s. The preheating temperatures were maintained between 350 and 400 °C to achieve steady-state combustion. Notably, the maximum combustion efficiency was 91%. Test 9 recorded the highest decomposition temperature of propellant in the catalyst bed (1422 °C). Interestingly, the combustion instability observed throughout this study was ≤0.5%. This study could assist in the further development of ADN-based propulsion systems. Full article

Ammonia dinitramide (ADN)-based liquid monopropellants are considered to be environmentally friendly alternatives to the toxic and carcinogenic hydrazine-based propellants. Hence, Space Solutions Co., Ltd. is developing a 1N ADN-based liquid monopropellant thruster by conducting a combustion performance in different types of reactors. Various parameters, such as preheating temperature and the size of thermal and catalyst beds, were examined. The results showed that the decomposition of the propellant in a Pt-LHA catalyst bed, which was used in the Type-1 reactor, resulted in insufficient combustion at low preheating temperatures. Furthermore, increasing the preheating temperature led to partial reaction of the propellant, but resulted in low combustion efficiency due to disintegration of the catalyst. However, when a thermal bed (STS ball) was used in addition to the catalyst bed (Pt-LHA) in the Type-2 and Type-3 reactors, the combustion efficiency was improved, with a minimal pressure drop of 0.2 bar. It was also confirmed that the catalyst was not damaged even after repeated operations. In conclusion, this study suggests that the propellant needs to vaporize before decomposing on the catalyst bed to achieve optimal combustion efficiency. Full article

The spray atomization of an injector significantly influences the performance and working life span of a bipropellant thruster of a spacecraft. Deep space exploration requires the thruster to be able to operate reliably at a low temperature range from −40 °C to 0 °C, so the effect of low temperature conditions on the atomization characteristics of injector spray is motivated to be comprehensively investigated. To study the swirl atomization characteristics of MMH (methylhydrazine), which is more difficult to atomize than NTO (nitrogen tetroxide), numerical simulations were conducted, employing the methods of VOF (volume of fluid) and LES (large eddy simulation) under low temperature conditions. The physical model with a nozzle size of 0.5 mm and boundary conditions with a velocity inlet of 3.89 m/s both follow the actual operation of thrusters. The development of spray atomization at low temperatures was observed through parametric comparisons, such as spray velocity, liquid total surface area, droplet particle size distribution, spray cone angle and breakup distance. When the temperature decreased from 20 °C to −40 °C at the same condition of flowrate inlet, those atomization characteristics of MMH propellant vary following these rules: the spray ejection velocity of MMH is significantly reduced by 7.7%, and gas-liquid disturbance sequentially decreases; the liquid film development is more stable, with a negative influence on atomization quality, causing difficulties for primary and secondary breakup, so the total surface area of droplets also decreases by 6.4%; the spatial distribution characteristics, spray cone angle and breakup distance vary less than 5%. Therefore, the low temperature condition can directly lower the combustion efficiency of thrusters with obvious performance degradation, but there are no significant changes in the propellant mixing and liquid film cooling. It is concluded that the bipropellant thruster can reliably work at low temperatures around −40 °C for deep space probe operation. Full article

An experimental study was carried out on the ignition characteristics of the HAN/(Emim)(EtSO₄) (hydroxylammonium nitrate and 1-ethyl-3-methyl-imidazolium ethyl sulfate) dual-mode ionic liquid monopropellant in chemical propulsion mode in model thrusters. Firstly, a model thruster with a detachable convergent nozzle was designed and fabricated. Secondly, catalytic ignition experiments at different flow rates were carried out in atmosphere and in high chamber pressure environment, respectively, using a model thruster, with and without the convergent nozzle. During the catalytic ignition process, measurement methods such as thermocouple, pressure sensor, and flue gas analyzer were employed to obtain the temperature at different depths of the catalytic bed, the pressure of the combustion chamber, and the concentration variations of gaseous products CO, CO₂, CH₄, SO₂, NO, and NO₂. Then the three characteristic stages of water evaporation, HAN decomposition, and (Emim)(EtSO₄) combustion were analyzed at the initiation time, and the reaction characteristics in the process of the catalytic ignition were analyzed. In addition, the composition and concentration of the combustion products at equilibrium were theoretically calculated. The effects of temperature and pressure on the concentrations of five main gaseous products were studied. Finally, the exhaust gas of the three groups of catalytic ignition experiments under different pressure environments was separately collected and measured with gas chromatography (GC) when the experiments approached equilibrium, the result of which roughly agrees with the theoretical calculations. These results are of great significance for exploring the chemical propulsion of the dual-mode ionic liquid propellant and understanding its physical catalytic combustion mechanisms. Full article

Green propellants are currently considered as enabling technology that is revolutionizing the development of high-performance space propulsion, especially for small-sized spacecraft. Modern space missions, either in LEO or interplanetary, require relatively high-thrust and impulsive capabilities to provide better control on the spacecraft, and to overcome the growing challenges, particularly related to overcrowded LEOs, and to modern space application orbital maneuver requirements. Green monopropellants are gaining momentum in the design and development of small and modular liquid propulsion systems, especially for CubeSats, due to their favorable thermophysical properties and relatively high performance when compared to gaseous propellants, and perhaps simpler management when compared to bipropellants. Accordingly, a novel high-thrust modular impulsive green monopropellant propulsion system with a micro electric pump feed cycle is proposed. MIMPS-G500mN is designed to be capable of delivering 0.5 N thrust and offers theoretical total impulse I_tot from 850 to 1350 N s per 1U and >3000 N s per 2U depending on the burnt monopropellant, which makes it a candidate for various LEO satellites as well as future Moon missions. Green monopropellant ASCENT (formerly AF-M315E), as well as HAN and ADN-based alternatives (i.e., HNP225 and LMP-103S) were proposed in the preliminary design and system analysis. The article will present state-of-the-art green monopropellants in the (EIL) Energetic Ionic Liquid class and a trade-off study for proposed propellants. System analysis and design of MIMPS-G500mN will be discussed in detail, and the article will conclude with a market survey on small satellites green monopropellant propulsion systems and commercial off-the-shelf thrusters. Full article

Ionic liquid electrospray thrusters are capable of producing microNewton precision thrust at a high thrust–power ratio but have yet to demonstrate lifetimes that are suitable for most missions. Accumulation of propellant on the extractor and accelerator grids is thought to be the most significant life-limiting mechanism. In this study, we developed a life model to examine the effects of design features, operating conditions, and emission properties on the porous accelerator grid saturation time of a thruster operating in droplet emission mode. Characterizing a range of geometries and operating conditions revealed that modifying grid aperture radius and grid spacing by 3–7% can significantly improve thruster lifetime by 200–400%, though a need for explicit mass flux measurement was highlighted. Tolerance analysis showed that misalignment can result in 20–50% lifetime reduction. In addition, examining the impact of electron backstreaming showed that increasing aperture radius produces a significant increase in backstreaming current compared to changing grid spacing. A study of accelerator grid bias voltages revealed that applying a reasonably strong accelerator grid potential (in the order of a kV) can minimize backstreaming current to negligible levels for a range of geometries. Full article

We report on the development of a software tool, the Electrospray Propulsion Engineering Toolkit (ESPET), that is currently being shared as a web application with the purpose to accelerate the development of electrospray thruster arrays for space propulsion. ESPET can be regarded as a database of microfluidic properties and electrohydrodynamic scaling models that are combined into a performance estimation tool. The multiscale model integrates experimental high-level physics characterization of microfluidic components in a full-scale electrospray propulsion (ESP) microfluidic network performance solution. ESPET takes an engineering model approach that breaks the ESP system down into multiple microfluidic components or domains that can be described by either analytical microfluidic or reduced order numerical solutions. ESPET can be divided into three parts: a central database of critical microfluidic properties, a microfluidic domain modeler, and a microfluidic network solver. Two options exist for the network solution, a detailed multi-domain solver and a QuickSolver designed for rapid design and testing of simple three-domain reservoir-feed-emitter arrays. The multi-domain network solver exploits the Hagen–Poiseuille/Ohm’s law analogy by using the publicly available SPICE (Simulation Program with Integrated Circuit Emphasis) electric circuit simulation software to solve the flow properties of the microfluidic network. Both the multi-domain and QuickSolver solutions offer Monte Carlo analysis of arrays based on user supplied tolerances on design parameters. Benchmarking demonstration examples are provided for experimental work in the literature, as well as recent experimental work conducted at Busek Co. The demonstration examples include ionic liquid propelled systems using active and passive capillary emitters, externally wetted emitter needles, and porous glass emitters, as well as a liquid metal system based on an externally wetted emitter needle. Full article

Liquid propellants are fast becoming attractive for pulsed plasma thrusters due to their high efficiency and low contamination issues. However, the complete plasma interaction and acceleration processes are still not very clear. Present paper develops a multi-layer numerical model for liquid propellant PPTs (pulsed plasma thrusters). The model is based on a quasi-steady flow assumption. The model proposes a possible acceleration mechanism for liquid-fed pulsed plasma thrusters and accurately predicts the propellant utilization capabilities and estimations for the fraction of propellant gas that is completely ionized and accelerated to high exit velocities. Validation of the numerical model and the assumptions on which the model is based on is achieved by comparing the experimental results and the simulation results for two different liquid-fed thrusters developed at the University of Tokyo. Simulation results shows that up-to $50\%$ of liquid propellant injected is completely ionized and accelerated to high exit velocities (>50 Km/s), whereas, neutral gas contribute to only $7\%$ of the total specific impulse and accelerated to low exit velocity (<4 Km/s). The model shows an accuracy up-to $92\%$ . Optimization methods are briefly discussed to ensure efficient propellant utilization and performance. The model acts as a tool to understand the background physics and to optimize the performance for liquid-fed PPTs. Full article