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Many plasma systems and technologies, such as Hall thrusters for spacecraft propulsion, exhibit complex underlying physics that affect the global operation. When characterizing such systems in an experiment, obtaining full spatiotemporal maps of the involved state variables can be, thus, highly informative. However, this goal is not practically realizable because of various experimental limitations, e.g., finite spatial resolution of the diagnostics and geometrical accessibility constraints. Therefore, having the capability to reconstruct the full high-dimensional states of plasma systems from low-dimensional time-history measurements is greatly desirable. Compressed sensing is a signal processing technique that can answer this crucial need. However, existing compressed sensing approaches have several limitations that restrict their effectiveness for complex physical systems like plasma technologies. These include the need for abundant sensor measurements and a principled sensor placement. In this paper, we demonstrate the capabilities of Shallow Recurrent Decoder (SHRED) architecture for compressed sensing. We show in several plasma test cases that SHRED can robustly infer full high-dimensional spatiotemporal state vectors of these systems (i.e., all macroscopic plasma properties) from minimal system information. This minimal information can consist of three finite time-history measurements of either local values of a plasma property or the global plasma properties (spatially averaged or performance parameters). An application of SHRED's inference capability in the numerical plasma simulation context is "super-resolution" enhancement. We will discuss this application by presenting how SHRED can effectively establish mappings between a low-resolution and a high-resolution simulation, recovering detailed spatial plasma features that are below the simulation's grid size. [ABSTRACT FROM AUTHOR]

Ionic liquid electrospray thrusters represent an alternative propulsion method for spacecraft to conventional plasma propulsion because they do not require plasma generation, which significantly increases the thrust efficiency. The porous emitter thruster has the advantages of simple propellant feeding and multi-site emissions, which miniaturize the thruster size and increase thrust. However, the multi-scale nature, that is, nano- to micrometer-sized menisci on the millimeter-size porous needle tip, makes modeling multi-site emissions difficult, and direct observation is also challenging. This paper proposes a simple model for multi-site emissions, which assumes that the ionic conductivity or ion transport in the porous media determines the ion-emission current. The conductivity was evaluated by comparing the experimental and numerical data based on the model. The results suggest that the ionic conductivity of the porous emitter is suppressed by the ion–pore wall friction stress. Additionally, the model indicates that the emission area expansion on the porous emitter creates the unique curve shape of the current vs voltage characteristics for multi-site emissions. [ABSTRACT FROM AUTHOR]

A multi-scale simulation framework for ion–solid interactions in plasma-exposed materials provides crucial insight into advancing fusion energy and space electric propulsion. Leveraging binary-collision approximation (BCA) simulations, the framework uniquely predicts sputter yields and analyzes material transport within volumetrically complex materials. This approach, grounded in the validated BCA code TRI3DYN, addresses key limitations in existing models by accurately capturing ion–solid interaction physics. A case study is presented, highlighting the framework's ability to replicate experimental sputter yield results, underscoring its reliability and potential for designing durable materials in harsh plasma environments. Insights into sputtering transport phenomenology mark a significant advancement in material optimization for improved resilience in plasma-facing applications. [ABSTRACT FROM AUTHOR]

Catalysts are widely used in research and in industrial applications to enable or to accelerate chemical reactions. One application is monopropellant thrusters for space propulsion systems, which use hydrogen peroxide (H2O2) as liquid propellant. The decomposition of liquid hydrogen peroxide into water and oxygen gas, which then finally generates the thrust, can be achieved using noble metal catalysts like platinum and iridium. In this study, iridium–platinum alloys were deposited onto ceramic pellets by a controlled magnetron sputtering process with different parameters. Selected parameters result in coated layers featuring an atomic structure of closed- and/or open-shell microstructures. The catalytic performance of these coated pellets was evaluated in laboratory experiments. The reactivity of sputtered iridium–platinum layers for the decomposition of H2O2 was found to be significantly higher with respect to layers of pure iridium coatings. This can be explained by the better reactivity of iridium–platinum alloys, combined with the active control of the surface morphology and the microstructure of the alloy coatings. [ABSTRACT FROM AUTHOR]

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

The oxide-rich lunar surface regolith can be used to extract the oxygen needed for the future of lunar exploration efforts as a consumable for life-support systems and spacecraft propulsion. Various techniques for the extraction of oxygen have been developed already, with solar vapor-phase pyrolysis shown to be a promising yet understudied approach. In contrast to other techniques, it requires only locally available resources, such as unbeneficiated regolith, sunlight, and vacuum in order to liberate oxygen and oxygen-bearing molecules. This study presents experimental work conducted in a purpose-built solar-vacuum furnace showing the evaporation of sodium and iron from a regolith simulant sample and their deposition on the crucible surface. This is matched by the thermochemical equilibrium modeling done in FactSage, which analyzes the process at varying pressures down to ultra-high vacuum. It highlights the need for precise temperature and pressure control, as well as the impact of regolith composition on oxygen dissociation for an efficient extraction of molecular oxygen. • The thermochemical model of the process shows molecular oxygen yields of up to 14.1%. • Low-Titanium Mare is the most promising composition for oxygen production. • Vapor-phase pyrolysis of regolith was demonstrated in a small solar furnace. [ABSTRACT FROM AUTHOR]

A diffractive solar sail is an elegant concept for a propellantless spacecraft propulsion system that uses a large, thin, lightweight surface covered with a metamaterial film to convert solar radiation pressure into a net propulsive acceleration. The latter can be used to perform a typical orbit transfer both in a heliocentric and in a planetocentric mission scenario. In this sense, the diffractive sail, proposed by Swartzlander a few years ago, can be considered a sort of evolution of the more conventional reflective solar sail, which generally uses a metallized film to reflect the incident photons, studied in the scientific literature starting from the pioneering works of Tsander and Tsiolkovsky in the first decades of the last century. In the context of a diffractive sail, the use of a metamaterial film with a Littrow transmission grating allows for the propulsive acceleration magnitude to be reduced to zero (and then, the spacecraft to be inserted in a coasting arc during the transfer) without resorting to a sail attitude that is almost edgewise to the Sun, as in the case of a classical reflective solar sail. The aim of this work is to study the optimal (i.e., the rapid) transfer performance of a spacecraft propelled by a diffractive sail with a Littrow transmission grating (DSLT) in a three-dimensional heliocentric mission scenario, in which the space vehicle transfers between two assigned Keplerian orbits. Accordingly, this paper extends and generalizes the results recently obtained by the author in the context of a simplified, two-dimensional, heliocentric mission scenario. In particular, this work illustrates an analytical model of the thrust vector that can be used to study the performance of a DSLT-based spacecraft in a three-dimensional optimization context. The simplified thrust model is employed to simulate the rapid transfer in a set of heliocentric mission scenarios as a typical interplanetary transfer toward a terrestrial planet and a rendezvous with a periodic comet. [ABSTRACT FROM AUTHOR]

The integration of superconductors into space propulsion has emerged as a promising avenue. Superconductors have demonstrated significant effectiveness in space propulsion devices in terms of performance. This study introduces a low-power High-Temperature superconducting Applied Field MPD thruster, utilizing High-Temperature superconductors to propel plasma out of the thruster to generate thrust. Replacing traditional copper magnets with High-Temperature superconducting magnets has not only led to substantial improvements but has also addressed challenges such as the increased weight of traditional copper coils, the energy-intensive nature of copper magnets consuming high power, and the extremely low temperatures required by conventional superconductors, necessitating the use of helium cooling tanks, which can add to the overall weight of the thruster. The research combines numerical simulations with detailed experimental validation to evaluate performance. As predicted by the simulation model, experimental results demonstrate a thrust of 220 mN at 15 mg/s and 15.2 kW, and 180 mN at 5 mg/s and 14.8 kW, with a specific impulse of 2760s, while maintaining a discharge current of 200 A. Additionally, the discharge voltage in the experimental setup increases with magnetic field strength but decreases with mass flow rate. Plasma fall voltage also increases with both the applied magnet field strength set by the High-Temperature superconducting magnet and the mass flow rate. Meanwhile, the fractional electrode fall voltage decreases with the applied field. Discharge voltage, plasma fall voltage, and electrode fall voltage are analyzed as factors influencing the increase or decrease in thrust, specific impulse, and efficiency of the High-Temperature superconducting-enhanced Applied Field Magnetoplasmadynamic thruster. Furthermore, the study concludes with a comprehensive examination of the High-Temperature superconducting-enhanced Applied Field Magnetoplasmadynamic thruster, offering valuable insights for researchers. Notably, the achieved efficiencies for Applied field Magnetoplasmadynamic thruster at low input power surpass those of the same configuration with traditional copper magnet, highlighting the potential benefits of utilizing High Temperature superconducting magnets for enhanced performance. • MHD simulation was considered. • PDE's were solved uisng Finite Element method. • High temperature superconducting Applied field MPD thrusters. • Low power HTS_AFMPDT. • Impulse of 2760s -with 14 % efficiency is achived. [ABSTRACT FROM AUTHOR]

Thrust generation through pulse-laser ablation has been proposed for use in propulsion systems in space. However, to achieve a high thrust performance, the energy efficiency must be improved and it is important to determine the behavior of the ablation plume to identify where energy losses occur. In this study, the ion-current distribution of a plume with aluminum ablation was measured. A neodymium-doped yttrium aluminum garnet (Nd:YAG) laser with a wavelength and pulse width of 1064 nm and 9 ± 2 ns, respectively, was used for the ablation. Experiments were conducted at three different fluences: 10, 15, and 20 J/cm2. The fluence was varied in two ways: changing the beam-spot diameter while maintaining the single-pulse energy and changing the single-pulse energy while maintaining the beam-spot diameter. The ion current was measured at various angles using Faraday probes. The generated impulse and reduction mass were also measured. By combining and analyzing the results, the divergence angle of momentum, mean valence of ablated aluminum, and energy efficiencies of several processes were estimated. • Aluminum pulse-laser ablation plume was investigated aiming use as space propulsion. • Nd:YAG laser with 1064 nm wavelength and 9 ± 2 ns pulse width used. • The ion current was measured at various angles using Faraday probes. • The generated impulse and reduction mass were also measured. • Divergence angle, mean ion valence, and energy efficiencies were estimated. [ABSTRACT FROM AUTHOR]

Multimode spacecraft propulsion has the potential to greatly increase the maneuvering capability of spacecraft in comparison to traditional architectures. This technology combines two or more propulsive modes (e.g., chemical and electric) into a single system with a single propellant. Trajectory design techniques for spacecraft with this capability, however, are presently limited and typically require manual selection of the burn sequence. In this study, indirect optimal control formulations with automatic mode selection are developed and applied for the first time to multimode spacecraft with two modes of propulsion. Minimum-fuel transfers are solved using polar coordinates as well as using Modified Equinoctial Elements with perturbations. Propellant-constrained minimum-time problems are also solved for the first time using a penalty function approach. An interior-point constraint formulation is also provided. Sample transfers are developed for each coordinate choice and optimization objective and are compared to trajectories that use a single mode of propulsion. The results demonstrate viability of the proposed techniques and show that the multimode approach can reduce the time-of-flight in comparison to a low-thrust only trajectory while providing mass savings over high-thrust only solutions. • Indirect optimal control methods are used to solve multimode mission design problems. • Minimum-fuel and propellant-constrained minimum-time problems are solved. • The techniques extend immediately to hybrid propulsion. [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]

This paper outlines the development phases of a wave-driven Helicon Plasma Thruster for cutting-edge Low Earth Orbit (LEO) constellations. The two-stage ambipolar electric propulsion (EP) system combines the efficient ionization of an ultra-compact helicon reactor with plasma acceleration based on an ambipolar electric field provided by a magnetic nozzle. This paper reveals maturation challenges associated with an emerging EP system in the hundreds-watt class, followed by outlook strategies. A 3 cm diameter helicon reactor was operated using argon gas under a time-modulated RF power envelope ranging from 250 W to 500 W with a fixed magnetic field strength of 400 G. Magnetically enhanced inductively coupled plasma reactor characteristics based on half-wavelength right helical and Nagoya Type III antennas under capacitive (E-mode), inductive (W-mode), and wave coupling (W-mode) were systematically investigated based on Optical Emission Spectroscopy. The operation characteristics of a wave-heated reactor based on helicon configuration were investigated as a function of different operating parameters. This work demonstrates the ability of two-stage HPT using a compact helicon reactor and a cusped magnetic field to outperform today's LEO spacecraft propulsion. [ABSTRACT FROM AUTHOR]

The development of modern reusable launchers, such as the Themis project with its LOX/LCH4 Prometheus engine, CALLISTO—a reusable VTVL-launcher first-stage demonstrator with a LOX/LH2 RSR2 engine, and SpaceX's Falcon 9 with its Merlin 1D engine, underscores the need for advanced control algorithms to ensure reliable engine operation. The multi-restart capability of these engines imposes additional requirements for throttling, necessitating an extended controller-validity domain to safely achieve low thrust levels across various operating regimes. This capability also increases the risk of component failure, especially as engine parameters evolve with mission profiles. To address this, our study evaluates the dynamic reliability of reusable rocket engines (RREs) and their subcomponents under different failure modes using multi-physics system-level modelling and simulation, with a particular focus on turbopump components. Transient condition modelling and performance analysis, conducted using EcosimPro-ESPSS software (version 6.4.34), revealed that turbopump components maintain high reliability under nominal conditions, with turbine blades demonstrating significant fatigue life even under varying thermal and mechanical loads. Additionally, the proposed predictive model estimates the remaining useful life of critical components, offering valuable insights for improving the longevity and reliability of turbopumps in reusable rocket engines. This study employs deterministic, thermally dependent structural simulations, with key control objectives including end-state tracking of combustion chamber pressure and mixture ratios and the verification of operational constraints, exemplified by the LUMEN demonstrator engine and the LE-5B-2 engine class. [ABSTRACT FROM AUTHOR]

Porous ionic electrospray emitters have received significant interest for space propulsion due to their performance and operational simplicity. We have developed a diffusion equation for describing the transient flow response in a porous electrospray emitter, which allows for the prediction of the settling time for flow in the porous emitter. This equation accounts for both the change in liquid storage at exposed pores on the emitter with pressure and viscous diffusion through Darcy's law. Transient flow solutions are provided for the most common emitter topologies: pillar, cone, and wedge. Transient flow solutions describe the settling time and magnitude of current overshoot from porous electrosprays, while providing useful guidelines for reducing transient response time through emitter design. Comparing diffusion of pressure to the onset delay model for electrospray emission shows that diffusion is most relevant at higher voltages and when a porous reservoir is used. Accounting for multiple emission sites on the wedge geometry shows that emission sites settle in proportion to emission site spacing to the power − 1.74. Article Highlights: The transient response of porous electrosprays is affected by fluid storage in exposed pores. Pressure diffuses into a porous electrospray through depletion of fluid in exposed pores. Diffusion of pressure complements the existing onset delay model for porous electrosprays. [ABSTRACT FROM AUTHOR]

A disruptive Electric Propulsion system is proposed for next-generation Low-Earth-Orbit (LEO) small satellite constellations, utilizing an RF-powered Helicon Plasma Thruster (HPT). This system is built around a Magnetically Enhanced Inductively Coupled Plasma (MEICP) reactor, which enables acceleration of quasi-neutral plasma through a magnetic nozzle. The MEICP reactor features an innovative design with a multi-dipole magnetic confinement system, generated by neodymium iron boron (NdFeB) permanent magnets, combined with an azimuthally asymmetric half-wavelength right (HWRH) antenna and a variable-section ionization chamber. The plasma reactor is followed by a solenoid-free magnetic nozzle (MN), which facilitates the formation of an ambipolar potential drop, enabling the conversion of electron thermal energy into ion beam energy. This study explores the impact of an inhomogeneous magnetic field on the heating mechanism of the HPT and highlights its multi-mode operation within a pulsed power range of 200 to 500 W of RF. The discharge state, characterized by high-energy electron-excited ions and low-energy excited neutral particles in the plasma plume, was analyzed using optical emission spectroscopy (OES). The experimental testing campaign, conducted under pulsed power excitation, reveals that, as RF input power increases, the MEICP reactor transitions from inductive (H-mode) to wave coupling (W-mode) discharge modes. Spectrograms, electron temperature, and plasma density measurements were obtained for the Helicon Plasma Thruster within its operational envelope. Based on OES data, the ideal specific impulse was estimated to exceed 1000 s, highlighting the significant potential of this technology for future LEO/VLEO space missions. [ABSTRACT FROM AUTHOR]

Hall thrusters are one of the most successful and prevalent electric propulsion systems for spacecraft in use today. However, they are also complex devices and their unique E × B configuration makes modeling of the underlying plasma discharge challenging. In this work, a steady-state model of a Hall thruster is developed and a complete analytical solution presented that is shown to be in reasonable agreement with experimental measurements. A characterization of the discharge shows that the peak plasma density and ionization rate nearly coincide and both occur upstream of the peak electric field. The peak locations also shift as the thruster operating conditions are varied. Three key similarity parameters emerge that govern the plasma discharge and which are connected via a thruster current–voltage relation: a normalized discharge current, a normalized discharge voltage, and an amalgamated parameter, α ¯ , that contains all system geometric and magnetic field information. For a given normalized discharge voltage, the similarity parameter α ¯ must lie within a certain range to enable high thruster performance. When applied to a krypton thruster, the model shows that both the propellant mass flow rate and the magnetic field strength must be simultaneously adjusted to achieve similar efficiency to a xenon thruster (for the same thruster geometry, discharge voltage, and power level). [ABSTRACT FROM AUTHOR]

This study aims to determine the optimal design and ascent trajectory for a single-stage hybrid rocket in the context of a lunar ascent mission, with the objective of reaching a 100 km circular orbit on the Moon equatorial plane. A multi-disciplinary optimization problem is formulated, integrating flight design variables and launch vehicle design variables to achieve the best mission scenario. A zero-dimensional internal ballistics model is employed to estimate key parameters such as thrust, mass flow rate, and gross mass of the hybrid rocket, while a three-degree of freedom dynamical model is utilized to simulate the ascent trajectory. The resulting integrated optimization problem is solved using an in-house heuristic algorithm that combines particle swarm and genetic algorithms. Two propellant combinations, liquid oxygen/paraffin-wax and hydrogen peroxide (90%)/high-density polyethylene, are evaluated, considering both their performance in the integrated system and the feasibility of the proposed mission. After conducting a sensitivity analysis to investigate the influence of particle count in the optimization process, results for both propellant combinations are presented. While paraffin and oxygen provide the highest payload ratio, hydrogen peroxide and polyethylene entail a more feasible and flexible design for a lunar mission. • Multi-disciplinary optimization of single-stage hybrid rocket for lunar ascent. • In-house hybrid heuristic optimization algorithm. • Evaluation of LOX/Wax and H2O2(90%)/HDPE. • LOX and H2O2 offer similar performance in terms of payload ratio. [ABSTRACT FROM AUTHOR]

This paper extensively reviews hybrid rocket propulsion-related activities from combustion engine designs to launch tests. Starting with a brief review of rocket propulsion development history, a comparison among the three bi-propellant rocket propulsion approaches, and hybrid rocket engine design guidelines, a very thorough review related to hybrid rocket propulsion and its applications is presented in this paper. In addition to propellant choice, engine design also affects the hybrid rocket performance and, therefore, a variety of engine designs, considering, e.g., fuel geometry, swirl injection, ignition designs, and some innovative flow-channel designs are also explored. Furthermore, many fundamental studies on increasing hybrid rocket engine performances, such as regression rate enhancement, mixing enhancement, and combustion optimization, are also reviewed. Many problems that will be encountered for practical applications are also reviewed and discussed, including the O/F ratio shift, low-frequency instability, and scale-up methods. For hybrid rocket engine applications in the future, advanced capabilities and lightweight design of the hybrid rocket engine, such as throttling capability, thrust vectoring control concept, insulation materials, 3D-printing manufacturing technologies, and flight demonstrations, are also included. Finally, some active hybrid rocket research teams and their plans for flight activities are briefly introduced. [ABSTRACT FROM AUTHOR]

In this study, we investigate and propose an improved theoretical method to more accurately predict the performance of a field-emission electric propulsion (FEEP) thruster with its complex configuration. We identify critical flaws in the previous theoretical methods and derive corrected equations. Additionally, we define and implement the overall half angle of the Taylor cone to account for variations in the Taylor cone's half angle depending on the applied voltage. Next, we also establish an improved method of the electric filed simulation in a three-dimensional domain to accurately predict a trajectory of extracted ions and a resulting spatial beam distribution of the FEEP thruster by incorporating a configuration of the Taylor cone with the estimated overall half angle from the results of the present theoretical method. Through comparison with the experimental measurements, we found that the present improved methods for theoretical and electric field simulations can yield more accurate predictions than those of the previous methods, especially for higher V and Iem regimes, which correspond to the actual operating conditions of the FEEP thruster. Consequently, we anticipate that the proposed methods can enhance the reliability and efficiency of the design process by accurately predicting performance when developing the new FEEP thruster with its non-symmetric complex configuration to match specific thrust or spatial beam requirements. [ABSTRACT FROM AUTHOR]

• A micro pulsed plasma thruster was developed and tested. • Specific impulse reaches 1600 s and efficiency greater than 10 %. • The designed thruster has been used in Hangsheng-1 Satellite. Pulsed plasma thrusters appeal to the propulsive tasks on miniaturized satellites because of the low volume and high specific impulse. However, further increasing efficiency and reducing the size of pulsed plasma thrusters still require more effort. In this paper, the thruster's spark plug, plate structure, ignition system, charging system, electromagnetic compatibility system, and propellant have been particularly designed to maximize its performance and miniaturize its dimension. With the conducted designs, the whole mass of the micro-pulsed plasma thruster is 450 g, and its dimension is 90 mm * 90 mm * 35 mm. According to the experimental results, the specific impulse reached 1600 s, and the efficiency was more significant than 9 %. Thanks to its excellent performance, we have used this thruster for attitude control and drag compensation tasks on the Hang-sheng-1 Satellite. [ABSTRACT FROM AUTHOR]

Nuclear thermal propulsion is an evolving technology that can be utilized for long-distance space travel. This technology yields the advantage of a high thrust and specific impulse, but requires an examination of the potential design adjustments necessary to enhance its feasibility. The development of nuclear thermal propulsion requires a comprehensive understanding of the system-level behavior during transient and steady-state operation. This paper extends our previous research by including the proper handling of turbomachinery with multi-channel thermal hydraulic simulations only for steady-state solutions. The system-level approach presented here enables the treatment of the turbopump components through non-dimensional analysis that eliminates the assumption of constant efficiencies. All the other components within the system (e.g., reflector and core) can be discretized to multiple channels and layers, in which the full thermal hydraulic solution is established. The approach chosen here enables the realistic modeling of the propellant flow within the expander cycle by capturing the pressure losses, mass flow rate splits, and enthalpy gain for various operational conditions. The verification of the package is completed through point comparisons of previous investigations into similar system designs. Furthermore, sensitivity studies are used to benchmark the capabilities of the package and investigate solution variations due to the perturbation of operational conditions and regimes. The sensitivity studies performed here are important to capture variation in flow characteristics (e.g., temperature, pressure, mass flow rates) for different design objectives such as the thrust and specific impulse. This work demonstrates that system-level simulations lacking multi-channel capability and proper turbomachinery treatment may yield higher uncertainties in understanding the engine's response and characteristics to changing various requirements. This is extremely important when screening the design space of such propulsion systems and when transient simulations are required. [ABSTRACT FROM AUTHOR]

Centrifugally confined Direct Fusion Drive propulsion has the capability of greatly advancing humanity's ability to traverse and explore the solar system. p-11B is a promising fusion fuel to use in such a system because the reaction is aneutronic. However, a large amount of bremsstrahlung radiation is produced and must be converted into electricity to power the electric field used to help confine the fusion plasma. This work details an idealized analytical model to determine the viability of using a thermionic energy converter for this purpose. The model finds a maximum power density of 1 0 5 W/m 2 with a Carnot efficiency of 30%. The required electric and magnetic fields to produce net positive fusion power are > 350 MV/m and ∼ 30 T respectively. This data is then used to discuss the feasibility of using p-11B as a fusion fuel for Direct Fusion Drive. • A TEC transforms up to 1 0 5 W/m 2 of radiation into electricity from a p-11B fusion plasma • The p-11B fusion fuel requires an electric field > 350 MV/m and a magnetic field ∼ 30 T [ABSTRACT FROM AUTHOR]

• Commercial-of-the shelf (COTS) space qualified solenoid valves for small satellite missions are financially unviable. • Thrusters control spacecraft attitude, the response time, and parameters of the solenoid valve is critical in mission design. • A mathematical model was developed to find the optimized geometrical configuration of the solenoid valve. • This paper presents a practical solution for designing custom solenoid valves. • The developed mathematical model validated using COTS space qualified solenoid valves. Small satellites and spacecrafts use propulsion systems to stabilize their attitude, control orientation and deorbit it at the end of the mission life. The flow and thrust of the liquid propulsion system are controlled by the control valve, typically operated by solenoid. According to recent literature, space-qualified solenoid valves are expensive and financially non viable, especially for small satellite missions. In this paper, a detailed methodology is presented to design the solenoid valve dedicated for small satellites. This study includes design criteria focusing on geometrical configuration. The dynamic, and electro-magnetic models are used to characterize the solenoid valve. Dynamic modelling helps to determine the response time of valve, affecting the thruster minimum impulse bit. The physical dimensions of the designed solenoid valve are smaller than those available in commercial off-the shelf (COTS) solenoid valves. However, the efficacy of the proposed design is proven effective with the COTS space grade valve specifications, and it shows that the proposed valve consumes less power than the conventional valves. Finally, it is worth noting that despite their modest and compact physical dimensions, the proposed valve are ideal for small satellite missions. [ABSTRACT FROM AUTHOR]

This paper explores the wide-ranging topography of micro-propulsion systems that have been flown in different small satellite missions. CubeSats, known for their compact size and affordability, have gained popularity in the realm of space exploration. However, their limited propulsion capabilities have often been a constraint in achieving certain mission objectives. In response to this challenge, space propulsion experts have developed a wide spectrum of miniaturized propulsion systems tailored to CubeSats, each offering distinct advantages. This literature review provides a comprehensive analysis of these micro-propulsion systems, categorizing them into distinct families based on their primary energy sources. The review provides informative graphs illustrating propulsion performance metrics, serving as beneficial resources for mission planners and satellite designers when selecting the most suitable propulsion system for a particular mission requirement. [ABSTRACT FROM AUTHOR]

The tenth Space Flights Safety Symposium discussed the problems of Space flight safety, protective measures for space structures and the future strategies aimed at conducting safe and peaceful Space missions. The following topics were under discussion: protection of Space structures from space debris collisions and micrometeoroids; fire safety of Space vehicles; propulsion systems modeling and simulations; supercomputer predictive modeling for ensuring Space program safety; safety at launch and during landing or splash down. • The tenth Space Flights Safety Symposium discussed the problems of Space flight safety. • Protection of Space structures from space debris collisions and micrometeoroids. • Fire safety of Space vehicles; propulsion systems modeling and simulations. • Supercomputer predictive modeling for ensuring Space program safety. • Safety at launch and during landing or splash down. [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]

Subgrid scale models for unresolved turbulent fluxes are investigated, with a focus on combustion for space propulsion applications. An extension to the gradient model is proposed, introducing a dependency on the local burning regimen. The dynamic behaviors of the model's coefficients are investigated, and scaling laws are studied. The discussed models are validated using a DNS database of a high-pressure, turbulent, fuel-rich methane–oxygen diffusion flame. The operating point and turbulence characteristics are selected to resemble those of modern combustors for space propulsion applications to support the future usage of the devised model in this context. [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]

The demand for space propulsion systems is increasing due to the rising number of satellite launches. Electric thrusters gain considerable importance as highly efficient systems in space. As the thrust generation process involves electrically charged particles in motion, ion and plasma thrusters can cause electromagnetic noise, which may interfere with satellite electronics or radio communication. Our objectives are to quantify such emissions and to better understand their origin. We use two facilities to achieve this goal: a semi-anechoic chamber (SAC) with a dedicated vacuum chamber and a mode-tuned vacuum chamber, also known as a reverberation chamber (RVC). Here, we conduct measurements in both facilities on an electron-cyclotron resonance (ECR) thruster with a magnetic nozzle in order to assess its electromagnetic compatibility (EMC). This thruster emits a quasi-neutral plasma plume without requiring an additional neutralizer, thus, is especially suitable for conducting our studies. The thruster in operation acts as an electromagnetic noise source contributing to different frequency ranges. In addition to the excitation frequency band, the data reveals emissions occurring approximately one decade lower in frequency. These emissions depend on propellant mass flow, background pressure and excitation power, i.e., vary with the electron density of the plasma generated inside the ECR thruster. The thruster was characterized following primarily MIL-STD-461G and further ECSS-E-ST-20-07C. [ABSTRACT FROM AUTHOR]

Laser Thomson scattering (LTS) is a measurement technique that can determine electron velocity distribution functions in plasma systems. However, accurately inferring quantities of interest from an LTS signal requires the selection of a plasma physics submodel, and comprehensive uncertainty quantification (UQ) is needed to interpret the results. Automated model selection, parameter estimation, and UQ are particularly challenging for low-density, low-temperature, potentially non-Maxwellian plasmas like those created in space electric propulsion devices. This paper applies Bayesian inference and model selection to a Raman-calibrated LTS diagnostic in the context of such plasmas. Synthetic data are used to explore the performance of the method across signal-to-noise ratios and model fidelity regimes. Plasmas with Maxwellian and non-Maxwellian velocity distributions are well characterized using priors that span a range of accuracy and specificity. The model selection framework is shown to accurately detect the type of plasmas generating the electron velocity distribution submodel for signal-to-noise ratios greater than around 5. In addition, the Bayesian framework validates the widespread use of 95% confidence intervals from least-squares inversion as a conservative estimate of the uncertainty bounds. However, epistemic posterior correlations between the variables diverge between least-squares and Bayesian estimates as the number of variable parameters increases. This divergence demonstrates the need for Bayesian inference in cases where accurate correlations between electron parameters are necessary. Bayesian model selection is then applied to experimental Thomson scattering data collected in a nanosecond pulsed plasma, generated with a discharge voltage of 5 and 10 kV at a neutral argon background pressure of 7 Torr-Ar. The Bayesian maximum a posteriori estimates of the electron temperature and number density are 1.98 and 2.38 eV and 2.6 × 1018 and 2.72 × 1018 m−3, using the Maxwellian and Druyvesteyn submodels, respectively. Furthermore, for this dataset, the model selection criterion indicates strong support for the Maxwellian distribution at 10 kV discharge voltage and no strong preference between Maxwellian and Druyvesteyn distributions at 5 kV. The logarithmic Bayes' factors for these cases are −35.76 and 1.07, respectively. [ABSTRACT FROM AUTHOR]

Ammonium dinitramide (ADN, [ NH 4 ] + [ N (NO 2) 2 ] - ) and its related propellants are promising high energy density materials for new-generation space propulsion. In order to ensure their safe utilization, it is of primary importance to assess the risk of accidental combustion events such as detonation. Thus, focusing on ADN and its related propellant composed of ADN, monomethylamine nitrate, and urea with weight percentages of 40:40:20 (AMU442), we have studied the properties and steady structure of detonation propagating in gaseous mixtures formed by their thermal decomposition. The AMU442-based mixture exhibits higher von Neumann and Chapman–Jouguet temperatures and pressures than the ADN-based mixture. The study of their steady detonation structure reveals that the gaseous species resulting from the decomposition of AMU442 have higher detonability than the ones resulting from the decomposition of ADN alone. This is in contrast to a previous study about the safety of these propellants in their original (solid or liquid) phase, i.e., AMU442 has lower sensitivity/reactivity to incident impact than ADN. Thermochemical analyses performed for both mixtures show that the decomposition of HNO 3 plays a dominant role for the energy consumption and initiation of the reaction by releasing both OH and NO 2 . For the ADN-based mixture, the reactions involving HN (NO 2) 2 and HNO 3 are the most sensitive, whereas for the AMU442-based mixture, the most sensitive reactions involve CH 3 NH 2 , CH 2 NH 2 , and HNO 3 . Reaction pathway diagrams emphasize the higher complexity of the chemical pathways for the AMU442-based mixture because of the presence of [ CH 3 NH 3 ] + [ NO 3 ] - and CH 4 N 2 O in the initial mixture. An uncertainty quantification study demonstrated that the calculated induction lengths exhibit an uncertainty on the order of 50%. [ABSTRACT FROM AUTHOR]

The Alternative Low-Power Hybrid Ion Engine (alphie) is a new technology for space propulsion based on plasma. Its distinct characteristic is the counterflow of charges (ions and electrons) passing through its two-grid system. This means that electrons coming from an external cathode are accelerated toward the ionization chamber, in which a neutral gas (typically Ar) is injected. The strong magnetic field therein confines these electrons, which ionize and exchange energy with the propellant gas. Thus, the operation of alphie is strongly affected by the electrons coming from the external cathode and their collisions with the neutral atoms. This work studies the counterflow employing a particle-in-cell simulation of ions and electrons passing through a single hole as a function of the electron cathode currents (Ice) and potential drops between grids (VAC). Transparency of the grid system to ions and electrons and the ion current extracted by the grid system are studied under sweeps of these two parameters. The number of ionization events by each high-energy electron entering the ionization chamber is evaluated using a physical model based on the gas density and the cross section for ionization. These new ions are then extracted by the same electric field that accelerates the electrons inward. Thus, simulations are self-consistent, since the ionizing electron flow from the external cathode drives the ion outflow at the exit section of the two-grid system. The electrical transparency of the two-grid system to ions and electrons, related to the axial charge currents, is also studied under sweeps of aforementioned operation parameters. This new way to deal with ionizations can be useful to study other plasma thrusters in which electrons for ionization come from an external cathode without modeling the complex structure of the ionization chamber. [ABSTRACT FROM AUTHOR]

Ionic-liquid ion sources produce beams of charged particles through evaporation and acceleration of ions and charged droplets from the surface of an ionic liquid. The composition of the emitted beam can impact the performance of ion sources for various applications such as focused beams for microfabrication and space propulsion. Numerical inference is considered for quantification of the beam composition of an ionic-liquid ion source through determining the current fraction of different species along with providing uncertainty in inferred values. An analysis of previously presented data demonstrates the ability to quantify the presence of ion clusters, including the distinct presence of heavy ion clusters such as heptamers. Quantification of beam composition will be an important technique for quantitative comparison of different time-of-flight data. [ABSTRACT FROM AUTHOR]

The quad confinement plasma source is a novel plasma device developed for space propulsion applications, whose core is an E × B discharge with open electron drift. The magnetic field is produced by independently powered electromagnets able to generate different magnetic field topologies with the ultimate aim of manipulating the ion flow field for achieving thrust vectoring. In this work, we map the ion velocity in the plasma ejected from the quad confinement thruster with different magnetic configurations using non-intrusive laser-induced fluorescence diagnostics. Measurements show a steep ion acceleration layer located 8 cm downstream the exit plane of the discharge channel, detached from any physical boundary of the plasma source. In this location, the ion velocity increases from 3 to 10 km/s within a 1 cm axial region. The ion acceleration profile has been characterized under multiple testing conditions in order to identify the influence of the magnetic field intensity and topology on this peculiar ion acceleration layer. [ABSTRACT FROM AUTHOR]

The advent of space exploration missions, especially those aimed at establishing a sustainable presence on the Moon and beyond, necessitates the development of efficient propulsion and mission planning techniques. This study presents a comprehensive analysis of chemical and electric propulsion systems for spacecraft, focusing on optimizing propellant distribution for missions involving transfers from Low-Earth Orbit (LEO) to Geostationary Orbit (GEO) and the Lunar surface. Using mathematical modeling and optimization algorithms, we calculate the delta-v requirements for key mission segments and determine the propellant mass required for each propulsion method. The results highlight the trade-offs between the high thrust of chemical propulsion and the high specific impulse of electric propulsion. An optimization model is developed to minimize the total propellant mass, considering a hybrid approach that leverages the advantages of both propulsion types. This research contributes to the field of aerospace engineering by providing insights into propulsion system selection and mission planning for future exploration missions to the Moon, Mars, and Venus. [ABSTRACT FROM AUTHOR]

Micro-cathode arc thrusters (μCATs) are space propulsion options suitable for microspacecraft, and have recently attracted much attention because of their low electrical power requirements and simple, compact propellant systems. The plasma ionization process, however, is not currently well understood. In this study, a μCAT prototype was designed with the use of different specific materials for each component, enabling the study of the ionization process from the emissive light of excited and ionized states from different elements. ICCD emission spectroscopy and scanning monochromator measurements are conducted inside the electrode channel of a μCAT with a coaxial configuration. The excited state atomic spectral results and ionization state spectral results give us a comprehensive understanding of the behavior of neutral particles and ions during the discharge process of the μCAT. The ionization sequence follows in the order of C+, Al+, Ti+, Cu+, and Ti2+. From the duration of the spectral lines, C I has the shortest duration (16.9 μs), which shows that the ablation of the conductive film is mainly concentrated during the first half of the discharge, while anode and cathode ionization occur over the entire discharge duration. Additionally, Al+ ionization is observed during the discharge process, showing that part of the discharge energy is expended in ionizing the thruster outer shell, resulting in efficiency loss. An ionization ratio of k is proposed in this work to represent the relative ionization of the conductive film and cathode material, aiding us in determining the optimal operation conditions for lifetime extension. • A μCAT prototype with different materials of each component is designed. • ICCD emission spectroscopy and scanning monochromator measurements are conducted inside a μCAT. • The impact of discharge energy and magnetic field strength on the ionization process is investigated. • An ionization ratio k is proposed to represent the relative ionization of the conductive film and cathode material. [ABSTRACT FROM AUTHOR]

The electric solar wind sail, or E-sail, is a propellantless interplanetary propulsion system concept. By deflecting solar wind particles off their original course, it can generate a propulsive effect with nothing more than an electric charge. The high-voltage charge is applied to one or multiple centrifugally deployed hair-thin tethers, around which an electrostatic sheath is created. Electron emitters are required to compensate for the electron current gathered by the tether. The electric sail can also be utilised in low Earth orbit, or LEO, when passing through the ionosphere, where it serves as a plasma brake for deorbiting—several missions have been dedicated to LEO demonstration. In this article, we propose the ESTCube-LuNa mission concept and the preliminary cubesat design to be launched into the Moon's orbit, where the solar wind is uninterrupted, except for the lunar wake and when the Moon is in the Earth's magnetosphere. This article introduces E-sail demonstration experiments and the preliminary payload design, along with E-sail thrust validation and environment characterisation methods, a cis-lunar cubesat platform solution and an early concept of operations. The proposed lunar nanospacecraft concept is designed without a deep space network, typically used for lunar and deep space operations. Instead, radio telescopes are being repurposed for communications and radio frequency ranging, and celestial optical navigation is developed for on-board orbit determination. [ABSTRACT FROM AUTHOR]

Hall thrusters are plasma-based devices that have established themselves as one of the most attractive and mature electric propulsion technologies for space applications. These devices often operate in a regime characterized by low frequency, large amplitude oscillations of the discharge current, which is commonly referred to as the 'breathing mode'. The intensity of these oscillations depends on the thruster's design and operating conditions and can reach values of the order of the average discharge current, posing issues for the thruster's performance and for coupling with the driving electronics. A 0D model of the thruster discharge was developed to investigate the core physical mechanisms leading to the onset and sustenance of the breathing mode. The model was found to be capable of reproducing oscillations with characteristics in line with those observed in the breathing mode. In this work, we extend the use of the 0D model to investigate the effect of the magnetic field intensity and of different propellants on the system stability. [ABSTRACT FROM AUTHOR]

Electric space propulsion is a technology that is used in a continuously increasing number of spacecrafts. The qualification of these propulsion systems has to run in ground-based test facilities which requires long testing times and powerful pumping systems. In these usually large test facilities, high pumping speeds are achieved with cryopumps. Cryopump operation is very expensive with respect to electrical energy and cooling water consumption. Therefore, being able to optimize pump shape, cold plate material, and pump placement in a chamber is beneficial. Pump design and tuned operating strategies can reduce costs and increase intervals between regeneration. Testing different pump configuration setups in a large facility is mostly prohibitive due to high costs and long testing times. Optimization via modelling is a better choice for design and also, later, for operation. Therefore, having a numerical model and proven guidelines at hand for optimization is very helpful. This paper describes a new model developed at DLR for the optimization of cryopump layout and operation. Model results are compared with cryopump operational and warm-up data. This validation is the basis for further optimization actions like multi-layer insulation layouts and pump cold plate upgrades, and helps in understanding and mitigating the detrimental effect of water condensates on the cryopump cold plates. [ABSTRACT FROM AUTHOR]

Magnetohydrodynamic (MHD) is one of the most promising novel propulsion technologies with the advantages of no pollution, high specific impulse, and high acceleration efficiency. As the carrier of this technology, the MHD accelerator has enormous potential for applications in hypersonic wind tunnels, supersonic ramjet engines, and deep space propulsion. In this study, a three-dimensional numerical simulation of an ideal Faraday magnetohydrodynamic (MHD) accelerator is conducted to assess the effect on performance with respect to applied potential and magnetic field intensity. The study is performed by employing a low magnetic Reynolds number MHD model coupled with a 7-component chemical reaction model to simplify the impact of real gas effects. The chemical reaction exhibits an increasing trend with rising applied potential and a decreasing trend with diminishing magnetic field strength. This variation influences the gas conductivity, subsequently affecting the velocity and thrust of the system. Specifically, at a magnetic field intensity of 2.0 T and an applied potential of 600 V, the accelerator exhibits maximum velocity and thrust growth rates of 18.6% and 59.8%, respectively. [ABSTRACT FROM AUTHOR]

Cathodic arcs (CA) are important across a wide range of applications including thin film deposition and solid fuel space propulsion. Understanding how cathode spots (CSs) move in magnetic fields has important implications for the deposition rates and thrust delivered as well as cathode/fuel usage efficiency. In this work, we investigated the dynamics of cathode spot (CS) motion in the presence of various axial magnetic fields (AMFs) in a high-current multi-spot CA system. Two photography techniques were used (i) long exposure and (ii) high-speed framing photography. The long-exposure images of cathode spots suggested that the cathode spots do not travel in spiral paths produced by Robson drift, as observed in previous vacuum interrupter system experiments. The reduced prevalence of Robson drift was found to be insufficiently explained considering only anodic plasma emission difference arising from the annular anode geometrical setup. High-speed framing images of cathode spots during discharges revealed a statistically significant reduction of the speed of cathode spot motion, as characterized by a decrease in K value as the applied AMF increases. [ABSTRACT FROM AUTHOR]

Liquefied gas storage and transportation, as well as space propulsion, are driving increasing interest in the cryogenic temperature deformation behaviour of 316L stainless steels. This was investigated here during tensile deformation at 15, 50 and 173 K. Complex transformation pathways, including one-step γ-austenite → α′-martensite, two steps γ → ε-martensite → α′ transformation as well as twinning and stacking fault-assisted γ → α′ transformation, are observed. γ grains with a (111) plane normal direction aligned 50–65°from the loading direction appear more likely to form the ε phase. Further, high-resolution transmission Kikuchi diffraction mapping revealed that the nucleation process of α′ can be assisted by ε and stacking faults at all cryogenic temperatures, whereas twins can also serve as sites for α′ nucleation when deformed at 173 K. For two-step transformation, separate lenticular α′ nucleate following Kurdjumov–Sachs orientation relationship (OR) within the shear band, once grown out of the shear band, Pitsch OR is preferable. As for one-step transformation, irregular oval α′ nucleates directly at γ grain boundaries with Nishiyama–Wassermann OR. These findings provide new insights into the correlation between the various transformation pathways and deformation mechanisms, as well as their improved performance at low temperatures. [ABSTRACT FROM AUTHOR]

In this paper, the injection characteristics of a liquid hydrogen propellant jet under subcritical and supercritical operating conditions are comparatively studied at the molecular scale using molecular dynamics methods in the context of spacecraft propulsion systems. To control the temperature and pressure within the environment and nozzle respectively, two sub-computational domains are set up for separate modeling. The sub-computational domain merging process eliminates the broken molecular bonds by using the computational domain reduction in the x-direction. The jet velocity, density and other physical property distributions at different temperatures and pressures were comparatively studied. It is found that the rate of change of the radial density of the jet under supercritical conditions is small. Based on this, the surface area and volume of the liquid phase region of the jet were extracted and calculated. Compared to result of subcritical case, the volume of the liquid hydrogen jet in the supercritical environment is reduced by 1.9%, while the surface area is increase by 6.7%. To explain the variation in the surface area due to molecular diffusion, the molecular stress results revealed that the stress in the hydrogen jet in supercritical environment is only 88.9% of that in subcritical environment. • Molecular dynamics is used to study the annular liquid hydrogen nano-scale jet. • The characteristic of jet in subcritical and supercritical ambient is studied. • The stress of jet in supercritical ambient is lower than subcritical result. • The jet in supercritical ambient tend to produce a larger surface area. • The subcritical jet has obvious stress stratification in atomic stress nephogram. [ABSTRACT FROM AUTHOR]

The work proposes the design of a novel internal combustion engine developed to benefit from the availability on board of cryogenic pressurized hydrogen and oxygen to serve a mechanical propulsion system. The rationale behind the design, as well as the results of simulations, are presented. The two-stroke engine uses direct injection of hydrogen, oxygen, and water, and downstream power turbines, to deliver power densities above 220 kW/L, zero-emission of nitrous oxides and carbon dioxide, and efficiency above 0.70. The novel engine could be an alternative to fuel cell electric propulsion systems, delivering in addition to better hydrogen fuel efficiency, also reduced weight and space of the complete propulsion system. • The availability of pressurized oxygen and hydrogen permit the design of high-efficiency engines. • A high-efficiency 2-stroke engine using oxy-hydrogen combustion is modeled. • The engine also feature water injection and two power turbines. • Computational results for this engine show the opportunity to achieve efficiency above 70%. • This is better than current fuel cells plus electric motors at a reduced weight and volume. [ABSTRACT FROM AUTHOR]

Space propulsion mainly uses reciprocating pumps that are fully integrated with small-scale applications. In this paper, the specific specifications and designing factors that should be met by rocket engine fuel pumps are demonstrated, and a comparative study is formed of the suitableness of all the necessary kinds of pumps to be used with rocket engines and their applications. Furthermore, it examines a piston pump intended to refuel the liquid fuel nozzles. Due to the performance of this pump, different parts and its performance have been evaluated. The design of the pump is discussed such that after completing the design steps according to the pump performance, a numerical solution of the stress and strain applied on the inner wall of the cylinder is performed according to its performance under internal pressure. Then, considering the consistency of the body material and the same pressure and heat applied to the cylinder and piston, a sample of the cylinder and piston is analyzed in finite element modeling technique and the modeling results obtained from the simulation are presented. [ABSTRACT FROM AUTHOR]

Direct numerical simulations (DNS) are conducted to analyze the effect of recess in the mixing and combustion performance of gaseous methane—oxygen injection systems. The recess length is varied from 0 to 2 injector diameters to analyze the sensitivity of various physical processes to this geometrical feature. It is found that the injection recess enhances the development of Kelvin–Helmholtz instabilities (KHI), due to the more consistent jet flow disposition at the injection plane, and the higher velocities. A Strouhal number Str R ≈ 0.22 , based on the injector diameter and on the average reacting jet velocity is found for all simulations. In addition, it is found that the enhancement of KHI acts as an injection of turbulent kinetic energy in the large scales, improving macroscale mixing and combustion performance. Finally, the normalized enhancements in combustion performance and averaged hydroxyl emissivity are found to be in good agreement with experimental results in a similar setup. [ABSTRACT FROM AUTHOR]

In this work, we investigate the properties of the plasma induced by focusing a high-power laser beam on an aluminum target that was cooled by a helium refrigerator from room temperature down to 20 K. Fast, streak photographs of the plasma were taken at different temperatures and laser energies. From the images obtained, position-vs-time plots were made for each experiment, and from them, the speed was calculated. Additionally, narrowband interference filters were employed to image the dynamics of ions and neutrals separately. It was found that the plasma plume has two distinct speeds: that of its center and that of the outer edge. For unfiltered images, the former has values within the interval 6.2 to 9.1 km/s, while the latter can reach speeds of the order of hundreds of km/s. It was found that the plume of a target cooled to 20 K has a length that is 8%–12% less than the corresponding size at room temperature. Chilling the target did not seem to affect significantly either the plume's speed of expansion or the size of the crater produced. Lower bounds were estimated for the momentum imparted to the ejecta and the specific impulse. The latter can reach a 920-s level, nearly twice as much the amount obtained with chemical rocket fuel. [ABSTRACT FROM AUTHOR]

During the 20th Century and within the first decade of the 21st, electric propulsion (EP) has played an important role in space activities, and its importance has not ceased to increase significantly in the last 15 years. The potential benefits of using EP on a bigger scale could be immense, reducing the cost of future space missions. The main hypothesis of this paper is that EP will play an increasingly important role in the system of international space relations in the coming decades with deep scientific, commercial, and geopolitical implications. To prove this, we examine the historical evolution, relevance, and future applications of such a propulsion system. Although there is extensive technical literature on EP, only a few works analyze electric space propulsion from an interdisciplinary perspective. On the contrary, much of the available information only provides an atomized, unipolar, and fragmented approach to the phenomenon. To fill this gap, this study presents an integrated analysis of the evolution, relevance, and trends of EP technologies, as well as their geopolitical implications in the space domain. The results obtained constitute a crucial contribution to the analysis of space propulsion from an interdisciplinary – socio-technological – perspective. [ABSTRACT FROM AUTHOR]

Applied field magnetoplasmadynamic thrusters (AF-MPDTs) are one of the potential thrusters for high power space propulsion. However, the high anode power deposition limits the improvement in the thruster's efficiency. According to previous research studies, the anode voltage drop is supposed to be the main factor of the anode power deposition in AF-MPDTs without considering the effect of the centrifugal force. Actually, the centrifugal force cannot be neglected when the magnetic field is high, where the anode voltage drop is not the only dominant factor of the anode power deposition. Due to the significant centrifugal force, there might be no obvious shortage of charge carriers near the anode surface in AF-MPDTs with propellant injection between the cathode and anode. The electrons will be compressed to the anode surface instead of the cathode. In order to investigate the reason for the anode power deposition in AF-MPDTs, a model considering the effect of the centrifugal force is established. The calculated anode power deposition shows good agreement with the corresponding experimental data. In addition, when the magnetic field is high, the viscous heating cannot be neglected, which is as significant as the anode sheath heating and Joule heating. Finally, in order to reduce the anode power deposition, a magnetic field distribution with a high value near the cathode and a low value near the anode is proposed. According to the evaluation by the model, the anode power deposition can be reduced by 30 % via the new magnetic field when the magnetic field exceeds 0.1 T. [ABSTRACT FROM AUTHOR]