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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]

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 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]

• 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]

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]

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]

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]

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]

Ionic liquids are regarded as promising propellants for green spacecraft propulsion. A well understanding of their fundamental properties is essential to accelerate the applications in practical propulsion. Therefore, the decomposition properties of an imidazole-based ionic liquid 1-ethyl-3-methyl imidazolium tetrafluoroborate ([EMIm][BF4]) were studied in this paper to help evaluate its chemical propulsion performance. Decomposition experiments were carried out with a high-temperature integrated thermal analyzer with heating rates of 5–20 °C/min. Results showed the decomposition process can be divided into four stages by three characteristic temperatures. The mass loss and heat absorption mainly occurred in the second and third stages. The activation energy of decomposition reactions was also obtained. Analyses indicated mechanisms in the first and second stages are multiple-step kinetics. But that of the third and fourth stages can be simplified as single-step kinetics. • Decomposition properties of [EMIm][BF4] were systematically studied. • Decomposition process was found to have four stages. • Activation energy of the decomposition reactions was obtained. • Decomposition process of [EMIm][BF4] should be described as multiple-step kinetics. [ABSTRACT FROM AUTHOR]

Cold/warm gas spacecraft propulsion systems generate thrust by accelerating a cold or moderately heated gas through a nozzle. Despite their simplicity compared with typical chemical or electric propulsion systems, the choice of propellant nonetheless represents an important consideration that affects the design, operation, and performance of such thrusters. Here we review different propellants that have been used or proposed for cold/warm gas propulsion systems, and we perform a study investigating almost 5000 other possible alternatives. Propellants include organic and inorganic substances that can be stored in solid, liquid, gaseous, and multi-phase states at ambient conditions, and which undergo a phase change to a pure gaseous state (where needed) by heating before being ejected from the thruster. The different propellants are assessed by considering important evaluation factors, such as propulsive performance, storage requirements and storage density, mission suitability, safety, and cost. While conventional thruster performance metrics like the specific impulse are low for some propellants, the design and operational advantages they offer, together with a higher total impulse mass density (i.e. the impulse per propulsion system wet mass), make them superior choices for many missions. [Display omitted] • 5000 substances investigated as possible propellants for cold/warm gas thrusters. • A new effective specific impulse metric is introduced to better assess performance. • Water and several other liquids give the highest effective specific impulse. • Xenon and krypton give the highest thrust-to-power ratio. • Several promising novel propellants (e.g. boric acid) have been identified. [ABSTRACT FROM AUTHOR]

Anomaly detection and section estimation methods using multiplexed fiber Bragg gratings (FBGs) have been developed to identify propellant supply anomalies rapidly and autonomously in spacecraft propulsion systems. FBGs were used to measure the external strain of the pipe due to internal pressure fluctuations in the pipe during the pressure response caused by a water hammer. The measured strain was substituted into the constructed conversion formula to obtain the differential pressure through the FBGs. As a result, compared with the differential pressure measured by a conventional pressure sensor at the same position, the error was 2% for the maximum pressure of the water hammer and 1% for the natural frequency, indicating high applicability as a differential pressure sensor in piping with FBG simply wrapped around the outside of the pipe. Since the peak intensity and frequency of the measured values in the frequency domain change during flow anomaly and bubble localization, the frequency response function (FRF) between response data of the section to be detected was calculated, and the least-squares error of the FRF was used as the degree of anomaly. As a result, flow anomalies and bubble localization became detectable and sectional estimation was possible. • Multiplexed FBGs enable highly accurate anomaly detection for spacecraft propulsion systems. • Simple and accurate pressure dynamic response measurement method realized by FBGs. • Anomaly evaluation and detection by frequency response function between normal and target data. • Estimation of anomaly occurrence section by round robin comparison of anomaly degree. [ABSTRACT FROM AUTHOR]

The National Aeronautics and Space Administration (NASA) has been involved in the development and maturation of metal additive manufacturing (AM) for space applications since the late 2000's. Several efforts have focused on the understanding of AM processes through material characterization and testing, standards development, component fabrication, and infusion into propulsion development and flight applications. NASA matured commonly used aerospace alloys from various alloy families (Nickel, Copper, Stainless and Steel, Aluminum, and Titanium-based) through detailed AM process and heat treatment characterization, in addition to mechanical and thermophysical testing. While these alloys are actively used in many propulsion applications, there is a need for ongoing AM optimized alloys using integrated computational materials engineering (ICME) and process development for high performance applications. The applications targeted are liquid rocket engines; advanced propulsion systems; and in-space propulsion with high heat fluxes, high pressure, and/or that use propellants that can degrade alloys (e.g., hydrogen). This paper highlights the characterization and physical properties of the more common AM alloys using laser powder bed fusion (L-PBF) and laser powder directed energy deposition (LP-DED) processes. Additionally, this paper discusses some of the ongoing novel alloy development and maturation using AM for use in these harsh environments, such as GRCop-42, GRCop-84, NASA HR-1, GRX-810, and C-103. The results from these processes demonstrated that AM could enable rapid development, and that optimized alloys could be developed using ICME, yielding higher performances. These alloys have undergone modeling, fundamental metallurgical evaluations, heat treatment studies, detailed microstructure characterization, and mechanical testing campaigns. This, combined with direct application-specific component fabrication and hot-fire testing, enabled the increase of the Technology Readiness Level (TRL) through high duty-cycle testing. A background and overview of these novel AM-enabled alloys and AM processing developments including metallurgical and mechanical property studies is presented here. The latest advancement in the parallel component development and hot-fire testing and future developments for these alloys is also discussed. • Additive manufacturing (AM) has enabled development of novel alloys for extreme propulsion environments. • NASA has matured alloys including GRCop-42, GRCop-84, NASA HR-1, GRX-810, C-103. • Characterization, material properties, and hot-fire test application data has been made available. [ABSTRACT FROM AUTHOR]

The electrohydrodynamic (EHD) actuator system, which produces thrust by ionizing a gas and accelerating it, is a novel propulsion system for near space vehicles. In this paper, an analytical model for EHD actuators, called EHD-AM, has been presented to estimate the thrust produced by the actuators in a wire-to-cylinder configuration. Experimental data from six previous studies have been used to validate the model and the mean absolute percentage error (MAPE) between each configuration and EHD-AM has been calculated under different environmental conditions. Based on the evaluations, the MAPE between EHD-AM and the experimental results for different pressures and wire-to-airfoil structures is 7.18% and 6.16%, respectively. The MAPE between EHD-AM and the experimental data measured by particle image velocimetry (PIV) method and with conventional experimental setups is 10.59% and 6.76%, respectively. Based on the EHD-AM, at a constant voltage, the relation between thrust and pressure has a local minimum and a local maximum point in the range of 0.05–1 atm. From 0.05 atm to 0.5 atm, as the pressure increases, the thrust decreases and then increases. As the pressure increases from 0.5 to 1 atm, the thrust decreases. In addition, a higher voltage can result in a higher pressure at which thrust is maximized. Furthermore, the generated thrust is monotonically increasing as the voltage amplitude increases. By implementation of EHD-AM thrust scale in a simple numerical model, the thrust generated by the actuator has been computed and simulated numerically. Initially, the numerical model was verified using experimental results from a previous study. The MAPE between the numerical model and the EHD-AM is 24.08%. Finally, when the interelectrode gap, the collector radius and the emitter radius are relatively large, the thrust of the actuator is predicted by EHD-AM and compared with the numerical results. EHD-AM can be used in the design and optimization of EHD actuators, and also for controller design and related analyses in a control system. • In EHD actuator, the relation between the thrust and the pressure is determined. • The relationship between corona current and voltage has been considered based on the space-charge-limited current relations. • In high and low mobility of ions, the thrust is obtained as a function of voltage. • Analytical model for corona discharge-based EHD plasma actuator is obtained by incorporating environmental conditions. [ABSTRACT FROM AUTHOR]

This paper explores the operational performance parameter range for inertial confinement fusion propulsion systems under the assumption of best case and worst case mission scenarios. This includes the spacecraft thrust, jet power and specific power. We show the derivation of the key driving equations and then simplify these with approximations for ease of domain space analysis. This includes a consideration of minimum and maximum values for jet efficiency, momentum weighting factor, nozzle divergence angle, fuel burn-up fraction, capsule mass and detonation pulse frequency. It is estimated that the range of parameter values may be of order 0.001 N–100 MN thrust, 1000 W–1,000 TW jet power, 0.01 W/kg–10 GW/kg specific power. Although this range includes values that are likely outside the realistic design space of applicability and to show this data from published designs is also examined. The analysis is considered for reaction isotopes of low mass hydrogen and helium nuclei only and does not take into account the possibility of propellant expellant augmentation, vehicle staging, design optimality or enhanced fuel reactions. • Attempt to examine inertial fusion propulsion theory from a broad perspective. • Derivation of driving equations for various inertial fusion propulsion parameters. • Analysis of min-max range inertial fusion propulsion performance space missions. • Analysis of inertial fusion propulsion performance data from concept designs. • Map of fusion propulsion parameters to derive the requirements for missions. [ABSTRACT FROM AUTHOR]

The knowledge of hazard classification of hydroxylammonium nitrate (HAN)-based liquid propellant is an essential subject in prevention of safe and transportation accidents. To improve hazard classification of the propellant, various tests were conducted on both material sample and packaged sample. Differential scanning calorimeter was utilized to study thermal decomposition parameters, resulting in the determination of high values for both thermal decomposition and explosion temperatures. Koenen and EIDS gap tests were used to evaluate the heat response under high confinement and the sensitivity of shock, respectively. The results indicated that the propellant was insensitive to shock wave, with a limiting diameter of 3.0 mm. The external fire test was conducted to elucidate the explosion characteristics of the packaged propellant. Pictorial information and temperature data were recorded using high-speed camera and infrared imager. The results revealed that a mass explosion occurred during the fire process. The combination of the perforated witness screen, long-range metallic projections and high-value heat flux led to the classification of the packaged HAN-based liquid propellant as Class 1.1 C, more stringent than the traditional Class 1.3 C. The study is expected to offer valuable insights for the safety of development, preservation, and conveyance of HAN-based liquid propellant. [ABSTRACT FROM AUTHOR]

In this study, a dual-axis impulse stand, having an elastic pivot with a precision of ±0.07 μNs, was newly developed, and the impulse induced by a laser pulse irradiation with a fluence of 10 J/cm2 was measured for incidence angles of 0°, 30°, 45°, and 60°. In all cases, the impulse vector was initially directed perpendicularly to the surface of the target. However, as the number of the laser pulse irradiations reached 200, only when the incidence angle was 60°, did the impulse direction shift to the laser-beam direction by approximately 10°. This impulse performance was consistent with the ablation crater morphology consisting in a periodic structure with 100 μm in width and 20 μm in depth, which was observed, through a laser microscope, only at this incidence angle. • A dual-axis impulse stand, having an elastic pivot with a precision of sub-μNs, was newly developed. • A laser-ablation impulse vector with oblique irradiation was investigated with the impulse stand. • When the incidence angle was 60°, the vector shifted to the laser-beam direction. [ABSTRACT FROM AUTHOR]

The problem of space debris must be addressed to avoid a cascading collision in the near future that would critically damage humanity's space assets and dependent terrestrial infrastructure. One method of meeting this challenge is to use a dedicated spacecraft to deorbit a series of prioritized debris. To avoid a large fuel requirement, electric spacecraft propulsion can be used, for which a multiple-rendezvous low-thrust trajectory is of interest. In this study, we develop such a trajectory for a predetermined subset of the Iridium 33 debris using a two-step procedure. First, the target rendezvous order is determined by using relevant distance metrics to approximate the transfer cost between debris objects. Next, the RQ-Law, a recently developed Lyapunov feedback control law for generating low-thrust rendezvous trajectories, is used to generate a trajectory that can take a spacecraft to each of the studied debris objects in the determined order. This method can be used to rapidly determine low-thrust multi-target multi-revolution rendezvous trajectories in the preliminary design stage for missions such as space debris removal and satellite servicing without requiring an initial guess. [ABSTRACT FROM AUTHOR]

The Hall-effect thruster is a prominent space propulsion system, providing a high specific impulse and thrust-to-power ratio. However, detrimental phenomena occurring in Hall thrusters, such as plasma instability and blowoff, are less understood. In this study, we employ a data-driven approach for analyzing the blowoff phenomenon of the Hall-effect propulsion system. From a 600-W class Hall thruster, we observe the oscillatory behavior of the anode current with a frequency of 34–38 kHz, which sometimes leads to plasma blowoff. We employ an output-only system identification framework that uses the noise-induced dynamics at this frequency, extracting linear and nonlinear coefficients of the low-order system model. The computed coefficients suggest that the plasma blowoff is characterized by large linear and small nonlinear parameters of the stochastic Van der Pol equation. From the obtained results, we conclude that this framework can be exploited either to identify a reliable operating condition with a low possibility of plasma blowoff or to diagnose the blowoff susceptibility of the thruster. This research constitutes the first application of stochastic system identification to the Hall-effect thruster, opening new possibilities for avoiding unwanted oscillations and blowoff phenomena in electric propulsion systems. • Discharge instability at 34–38 kHz leading to plasma blowoff of the Hall-effect thruster. • A data-driven approach using the system identification from the noise-induced dynamics of anode current oscillations. • Low-order modeling using stochastic Van der Pol oscillator and the time-domain analysis of plasma oscillation in the Hall thruster. [ABSTRACT FROM AUTHOR]

• Solar thermal propulsion system combined with thermal energy storage is proposed to overcome thrust failure in shadow area. • The whole process from light concentrating, heat storage, to trust generation is numerically investigated and verified by ground tests. • The time to complete heat storage is within the illumination time of low Earth orbit. • Thrust and specific impulse can achieve 2 N and 691 s respectively in solar eclipse region. Possessing relatively high specific impulse and moderate thrust levels, solar thermal propulsion (STP) is a promising candidate in spacecraft propulsion system. However, the traditional solar thermal propulsion system suffers from thrust failure in the shadow area, which seriously affects its applicability. In this paper, we investigate feasibility of regenerative solar thermal propulsion system (RSTP) incorporating thermal energy storage, which can effectively overcome unmatched synchronous working time and illumination time. A numerical model for RSTP considering the whole energy transfer process from light concentrating, heat storage, to thrust generation is built, which is verified by experiment measurements with relative errors less than 15 %. The result shows that the maximum time to complete heat storage is about 4000 s, which is within the illumination time for low Earth orbit. In the solar eclipse region, the thrust (F t) and the specific impulse (I sp) of the system increase with the propellant flow rate, which can reach about 2 N and 690 s, respectively. What's more, the system can operate for around 100 s continuously at the maximum thrust in the shadow area. This work provides alternative approaches for microsatellite propulsion with high specific impulse, high thrust, and continuous operation despite presence of solar eclipse. [ABSTRACT FROM AUTHOR]

• A capture mission scenario of a radially-propelled space-craft is analyzed. • The capture parabolic orbit is circularized with modulated radial thrust. • The problem is studied in an optimal framework to minimize the flight time. • The parameters defining the capture trajectories shape are identified. • A test case scenario involving a capture mission around Ceres is discussed. The aim of this paper is to analyze the performance of a radially-accelerated spacecraft in a capture mission scenario, in which a space vehicle transfers from a parabolic approaching trajectory of assigned semilatus rectum to a target circular orbit around a generic celestial body. The radial propulsive acceleration provided by the spacecraft propulsion system can be modulated within a suitable given range, and the radial thrust can be either inward- or outward-directed. The transfer mission performance is analyzed in an optimal framework, by minimizing the total flight time with an indirect approach. In this context, the paper proposes a semi-analytical method to reduce the computation time required to solve the optimal control problem. Using a proper description of the spacecraft dynamics in a dimensionless form, the paper reports a set of graphs and tables that allow the designer to obtain a rapid approximation of the optimal mission performance with an accuracy consistent with that of a preliminary trajectory design. Finally, the proposed approach is used to describe the minimum time capture trajectories in a mission scenario towards Ceres. [ABSTRACT FROM AUTHOR]

This work presents some of the first results from a recently developed fusion propulsion code which calculates the performance of an advanced spacecraft design. HeliosX is a system integrated programming design tool written in Fortran 95 for the purpose of calculating spacecraft mission profile and propulsion performance for inertial confinement fusion driven designs. This code uses the vehicle configuration input and then calculates the likely mission profile for a given destination target. The key capability under development is the inclusion of the fusion propulsion system and an adequate modelling of its likely energy outputs and how this integrates into the mission profile. The results of a comparison to a concept from the literature are firstly presented for an interstellar flyby probe and then perturbations of this design to include updates and revisions to its input parameters and expected performance. Although the code is still being developed the early results show good potential to enabling a numerical tool for conducting spacecraft performance assessment of fusion based propulsion systems. The preliminary results of modelling different missions are shown for an interstellar target in ∼ 37–100 years trip times and ∼ 0.05–0.12c cruise speeds in flyby and rendezvous modes for orbital insertion to a distance ∼ 4.3–5.9 l y. A description is given for how the baseline concepts are being evolved since the code has been developed in parallel with the design process which aims to design improved vehicle concepts. Ultimately the calculations shown in this paper are working towards the development of a design tool which can facilitate the proper analysis of a range of different spacecraft designs and place them on a consistent modelling basis so that feasibility assessments can be made. • First presentation of design tool that permits rigorous assessment of assumptions used in space propulsion calculations driven by an ICF propulsion engine. • Describes how the ICF capsule assumptions and performance outputs influences the integrated spacecraft design, systems and associated mission profile. • The code has been applied to a historical design as a form of numerical validation of its physics and engineering algorithms to include modern updates. • Perturbations of the design have been made as more engineering elements and algorithms have been added to the code to increase its design applicability. • A review has been conducted of the historical design to inform the generation of advanced baseline models and apply to a wider design suite. [ABSTRACT FROM AUTHOR]

In space propulsion hydrogen peroxide is a "green" solution that overcomes the problem of toxicity related to hydrazine-based systems. This work aims at designing, developing, and characterizing a monopropellant thruster for small satellites, based on hydrogen peroxide. The engine is intended for use on a 12-Unit CubeSat in Low Earth Orbit. Based on mission requirements a catalytic chamber and a nozzle were designed and a ground-test breadboard was manufactured. The prototype was assembled and tested on ground and measured engine performances are discussed highlighting their dependence on the nozzle throat diameter, that primarily affects the chamber pressure. Finally, effects of the mass flow rate on the system reactivity and steady state conditions are also investigated. • A 5 N HTP monopropellant thruster for cubesats was designed. • Tests were carried out to investigate the effect of throat diameter and flow rate. • Pressure affects residence time and increases system reactivity. • Higher flow rate makes system response faster, lower flow rate increases decomposition temperature. [ABSTRACT FROM AUTHOR]

Green propulsion is a key element of future Space propulsion because being environmentally friendly is not just a legal obligation but it can significantly reduce the hazards to people involved in the handling and thus also the overall lifecycle cost. This reduction is due to the easier handling of propellants compared to state-of-the-art toxic propellants based on Hydrazine. When it comes to a replacement of classical, toxic propellants for long term orbital missions current replacement candidates still have handling limitations because these propellants contain chemical energy. Water is the greenest of all possible propellants with no hazard potential because the energy is added in orbit via solar power. The Water Electrolysis Propulsion (WEP) is a system where gaseous oxygen and hydrogen are produced via electrolysis from pure water over a long period with low electric power and burnt with a high thrust and high Specific Impulse (ISP) when needed. Compared to classical Electric Propulsion Systems in this system the energy generation and usage are decoupled. With own R&T support, German (DLR) and European (ESA) funding this technology and the two key technology bricks (Electrolyser, stoichiometric thruster) have been further matured within ArianeGroup. This paper describes the latest development tests of the space grade electrolyser, the stoichiometric thruster and a coupled test series. [ABSTRACT FROM AUTHOR]

• Triglyme with a Cu(II)-based additive is proposed as a low-toxicity hypergolic fuel. • Hypergolic ignition was experimentally investigated using ≥ 70 wt% H 2 O 2. • The ignition delay time of triglyme-based fuels was improved by blending with an ionic liquid. • Theoretical performance of the proposed propellant is comparable to MMH/NTO. Low-toxicity hypergolic propellants have been a demanding research area in recent years for space propulsion, with the aim of replacing conventional toxic propellants. Therefore, the current research was focused on the development of low-toxicity hypergolic propellants for safe, cost-effective, and sustainable space exploration. In this study, triglyme was selected as a fuel component for the low-toxicity hypergolic propellant. Hypergolic performance of various concentrations (3 to 13 wt%) of monocyanotrihydroborate tetrakis(imidazole) copper(II) cyanotrihydroborate additive (Cu-P1) in triglyme fuel was experimentally investigated with 70, 90 and 95 wt% hydrogen peroxide (H 2 O 2). The new fuels, 13Cu-TriG and 13Cu-TriGIL, containing 13 wt% of Cu-P1 in triglyme and triglyme:1-ethyl-3-methyl imidazolium cyanoborohydride ([EMIM][BH 3 CN]) blend, respectively, were investigated. The hypergolic performance and properties of triglyme were improved by using [EMIM][BH 3 CN]. For example, 13Cu-TriG was not ignited with 70 wt% H 2 O 2 ; however, 13Cu-TriGIL exhibited an IDT of 16.5 ms. Furthermore, 13Cu-TriG and 13Cu-TriGIL revealed an IDT of 18.7 and 8.0 ms, respectively, with 95 wt% H 2 O 2. The physicochemical properties, such as density, viscosity, decomposition temperature, and freezing/melting temperature, of 13Cu-TriG and 13Cu-TriGIL, were also measured. 13Cu-TriGIL showed a density of 1.017 g/cm3 and a viscosity of 26.42 mPa·s. Additionally, 13Cu-TriG exhibited a freezing point of −45 °C. However, with the addition of [EMIM][BH 3 CN] in the fuel, the freezing temperature drops below −80 °C (13Cu-TriGIL). Furthermore, 13Cu-TriGIL with 95 wt% H 2 O 2 demonstrated theoretical specific impulse and density specific impulse of 2.1 % lower and 9.7 % higher than MMH/NTO, respectively. Overall, the use of ionic liquid to improve the hypergolic performance and physicochemical properties of triglyme-based fuels was the advantage of our research. [ABSTRACT FROM AUTHOR]

• MgH 2 was first applied as the H 2 and Mg sources for the dual-mode propulsion system of the spacecraft. • A novel MgH 2 hydrogen storage tank was proposed to adapt to zero gravity used for spacecraft. • The complete hydrogen desorption process of the MgH 2 tank at high backpressure was proved. • The parameters were optimized to achieve a high hydrogen desorption rate per unit thermal power. A dual-mode propulsion system stands as the important viable pathway for future multi-purpose deep space missions by the spacecraft, encompassing both nuclear thermal propulsion and nuclear electric propulsion. H 2 and Mg emerge as potential mediums for nuclear thermal propulsion and nuclear electric propulsion, respectively, with the possibility of utilizing a single material, magnesium hydride (MgH 2), to serve both purposes. However, how to release hydrogen completely and effectively in high hydrogen desorption pressure while retaining magnesium is the key problem. In this work, we proposed a novel MgH 2 storage tank for zero gravity environments and investigated its hydrogen desorption process at high pressures of 4 ∼ 8 MPa using simulation, which is crucial for the hydrogen supply of nuclear thermal propulsion. We examined the effects of hydrogen desorption pressure, inlet temperature, and flow velocity of argon (Ar) on the desorption process. The simulation results indicated complete decomposition of MgH 2 can be realized under 4 ∼ 8 MPa at 873 K. Moreover, the desorption pressure, inlet temperatures, and high flow velocities of Ar are suggested as 4 MPa, 773 ∼ 785 K, and 45 ∼ 100 m s−1 to achieve an average hydrogen desorption rate per unit thermal power of ≥ 20 mgH 2 s−1. This work marks the first confirmation that MgH 2 can decompose into H 2 and Mg under high hydrogen desorption pressure at suitable thermal power, paving the way for its application as the hydrogen and Mg storage media in the dual-mode propulsion system of spacecraft. [ABSTRACT FROM AUTHOR]

• The molybdenum material was used as an electrode to enhance the efficiency of HAN-based gel propellant fuel ignition. • The combustion wave structure of HAN-based gel propellant was divided. • Combustion mechanism of electric-thermal coupling on HAN-based gel propellant was established through the self-built combustion diagnosis system and instrumental analysis. Hydroxylammonium nitrate (HAN) serves as a monopropellant fuel, renowned for its low toxicity and high-performance in space propulsion systems. HAN-based gel propellant may offer additional advantages, including long-term stability and excellent throttling capabilities. To better understand the ignition performance and combustion characteristics of this gel propellant, we prepared HAN-based gel propellant, and monitored its decomposition products using thermogravimetry (TG), Fourier Transform Infrared Spectroscopy (FTIR) and mass spectrometry (MS). Electrode materials of Mo, Cu, Ti and SUS 304 were fabricated to investigated their impact on ignition performance. It was found that the Mo electrode exhibited the shortest ignition delay time (t i) compared to Cu, Ti, and SUS 304. Meanwhile, the role of surface roughness of Mo electrode on the ignition delaytime (t i) and the ignition energy (Q i) were also studied. Our results revealed that, as the voltage increased from 100 V to 200 V, the fine-surfaced Mo electrode demonstrated a lower sensitivity to changes in ti compared to the rough-surfaced electrode. Additionally, the Q i of the rough-surfaced electrode decreased from 43.5 J to 16.7 J, while the Q i of the fine-surfaced electrode decreased from 52.5 J to 17.6 J. Combining a thermocouple, the combustion wave structure of the gel propellant was divided into pre-heating zone, condensed phase, electro chemical reaction zone, microexplosion-gases mixing zone and flame zone. Finally, the combustion model of HAN-based gel propellant was inferred to explain its combustion mechanism. [ABSTRACT FROM AUTHOR]

• Constellation deployment problem generalization using finite propulsive transfers. • Introduction of algorithms for multi-objective constellation deployment optimization • Evaluation of impact of launchers, space environment and propulsion system adopted Constellation deployment using on-board propulsion is a cost-effective way for distributing multiple satellites into different orbital planes using a minimal number of launches. This paper focuses on the extension of previous studies based on analytical formulations of the problem into a more general numerically-based one, which can be applied to both electric and chemical propulsion systems. Additionally, the formulation is not restricted to in-plane maneuvers only, taking full advantage of the increased mobility capabilities of satellites equipped with high-efficiency propulsion systems. Two deployment methods are analyzed: the first is based on moving the satellites from the injection to the operational orbit sequentially with a properly timed delay between each maneuver so that the difference in nodal regression causes the accumulation of a given spacing; the second is built upon the parallel transfer of satellites into so called "drift" orbits which, once again, exploiting the differential nodal drift, can be used to accumulate orbital plane spacing until a required value is met, and the satellites can be finally moved into their operational orbits. A multi-objective optimization based on the high-fidelity numerical computation of maneuvers using a feedback control law coupled with a genetic algorithm is proposed, presenting a framework capable of generating optimal solutions with respect to total constellation propellant mass needs and deployment time. An example of application based on the FORMOSAT-3/COSMIC mission is presented. The results highlight the potential of the optimization routine which is intended to be applied systematically by mission designers to generate solutions for the deployment problem within the context of constellation design. Some important and un-explored properties of the problem, such as, launcher-related injection constraints, the selection of a suitable propulsion technology and the effects of atmospheric drag are analyzed. The results show the effect of these properties on the optimality of the solutions, and highlight the advantage of a mixed in-plane/out-of-plane maneuver approach. [ABSTRACT FROM AUTHOR]

In recent years, hydrogen peroxide (H 2 O 2) has been attracting attention as an alternative oxidizer to replace the traditional toxic hypergolic oxidizers. However, hypergolic solid fuel (HSF) candidates for H 2 O 2 are few in number and offer limited applications for space propulsions. In this work, various activated carbon (AC)-supported catalysts for a H 2 O 2 hybrid rocket were studied to improve the hypergolicity of HSF, thereby widening the selection range for hypergolic fuels. For this purpose, eleven H 2 O 2 catalyst candidates (Cr, Mn, Fe, Co, Ni, Cu, Ru, Pd, Pt, Pt–Ru, and Pb) were deposited onto AC supports at contents of 20 wt%. The nanostructure and metal particle size distribution of these prepared catalysts were investigated via transmission electron microscope imagery. The oxidation states of the supported catalysts were determined by X-ray diffraction, while the ignition characteristics of the pure catalysts and HSFs with catalyst additives were examined by conducting hypergolic drop tests using a 95 wt% H 2 O 2. The hypergolicity of the prepared catalysts was evaluated by measuring their reaction delay and ignition delay times. Among the studied catalysts, Pt/AC, Ru/AC, and Pt–Ru/AC exhibited the shortest reaction delay time of 0.3 ms. Moreover, the dependences of the reaction delay time on the presence of Na+ ions in Mn/AC and Mn concentration were established. Finally, the ignition delay time of an HSF containing 25 wt% ammonia borane and 75 wt% polyethylene decreased from 20.3 to 9.8 ms after the addition of 1 wt% Pd/AC. These results can provide a variety of additive options for improving hypergolicity of solid fuels with H 2 O 2. • Eleven activated carbon-supported catalysts are studied to improve the hypergolicity of solid fuels. • Reaction delay and ignition delay times of the fabricated catalysts are measured. • Ignition characteristics of hypergolic fuels are enhanced by catalysts addition. • The shortest reaction delay time of 0.3 ms is achieved. [ABSTRACT FROM AUTHOR]

This work focuses on a numerical suite targeted at cathodeless plasma thrusters. A Global Model is used for preliminary estimation of the propulsive performance. In the plasma source, the code 3D-VIRTUS solves self-consistently plasma transport with a fluid approach, including the electromagnetic wave propagation. Starfish and SPIS are used to handle the magnetic nozzle region via a fully kinetic Particle-in-Cell (PIC) model. Starfish solves the plasma dynamics in a two-dimensional axisymmetric domain and has been coupled to 3D-VIRTUS in order to estimate the propulsive performance. SPIS is a three-dimensional PIC code that has been used to simulate the mutual interactions between the satellite surfaces, the propulsion system plumes and the environmental plasma. Results have been benchmarked against thrust measurements performed on a low-power (50 W range) Helicon Plasma Thruster. • Cathodeless plasma thrusters for space propulsion application. • Numerical suite for both preliminary and detailed simulation of plasma thrusters. • Global model to preliminary estimate propulsive performance. • Fluid approach to simulate plasma source, Particle-In-Cell for the magnetic nozzle. • Experimental validation of the numerical results. [ABSTRACT FROM AUTHOR]

Here we present a detailed analysis of a new concept for beam propulsion that extends the propagation range through space by circumventing fundamental limitations on the beam divergence. Using a combination of spatially overlapped continuous laser and neutral particle beams, opto-mechanical forces and light refraction are simultaneously exploited to produce self-guided modes. The aim of the present work is to elucidate the relevant physical phenomena, their relation to the conditions for self-guiding, and practical considerations for the design parameters of such a propulsion system. Methods for computing the lowest order axisymmetric self-guided modes are presented along with the resulting spatial mode structure. Key non-dimensional parameters are derived, including the use of linear stability analysis to determine features of the dynamic interaction. Based on these findings, a beam parameter optimization framework is proposed and applied to a 100-year interstellar flyby mission to Proxima b. With 200 GW of combined laser and particle beam power, we find that a 0.3 kg spacecraft could be accelerated to 0. 044 c over a distance of 4. 37 × 1 0 6 km (0.029 AU) using a small 1 m 2 beam cross-section. Compared to an equivalent particle beam with no self-guiding, this represents an over 50 × increase in payload. Similar payload mass would be possible by laser propulsion with an approximately 1 0 4 m 2 area laser array. These results suggest that self-guided beam propulsion possesses unique attributes that could enable the acceleration of larger spacecraft to relativistic velocities and enhance scientific capabilities for future interstellar and interstellar-precursor missions. [ABSTRACT FROM AUTHOR]

In space propulsion applications, the development of new ceramic matrix composites with improved resistance to oxidation and ablation at high temperature is needed and ultra-high temperature ceramics-based ones appear the most suitable. Combination of both powder impregnation (ZrB 2 , C) and liquid silicon infiltration enabled manufacturing of UHTC based matrices in C f /C preforms with less than 10 vol% open porosity and various proportions and homogeneous distribution of C, ZrB 2 , SiC and Si. Oxidation behaviour was evaluated on composite structures using an oxyacetylene torch at temperatures higher than 2000 °C. Chemical analyses and microstructural observations before and after oxidation testing evidenced the protection ability of ZrB 2 -SiC-Si matrices thanks to the formation of multi-oxide scales which resisted even tested durations of 6 min and pointed the unharmful presence of residual 12 vol% silicon on the composite for use at high temperature under high gas flows. [ABSTRACT FROM AUTHOR]

Helium mass spectrometry leak detection technology requires that the leak detection medium of spacecraft test object is helium, but the general working medium of spacecraft electric propulsion system is xenon. The existing helium mass spectrometry leak detection technology can not realize the in-situ leak detection of all kinds of working media. The in situ mass spectrometry leak detection technology of spacecraft working medium based on mass spectrometry has been widely concerned for its advantages of direct leak detection of working medium and obtaining real leak rate, but the sampling and leak detection of mass spectrometry under atmospheric environment is one of its difficulties. In this paper, a leak detection model and test system of spacecraft working medium in situ mass spectrometry are established, and theoretical and experimental studies are carried out. The research results show that, The leak detection sensitivity of in situ working mass spectrometry is better than 1.0 × 10−8 Pa m3/s by shunt current limiting sampling method in atmospheric environment. The appropriate shunt pumping speed and shunt chamber pressure have important influence on in situ leak detection of working medium. When the pressure of capillary shunt chamber is higher than 550Pa, the test system has better stability and repeatability. Not only the background noise of the leak detection by mass spectrometry is reduced, the detection accuracy is improved, and the response time of the system is shortened. • This paper established a model based on flow limiting shunt sampling mass spectrometry. • This paper analyzed the correlation effect of the parameters of flow limiting shunt sampling. • The appropriate shunt pumping speed and shunt chamber pressure have important influence on in situ leak detection. • It is necessary to analyze and design the leak testing system from the theoretical and experimental aspects. [ABSTRACT FROM AUTHOR]

Load-following capability is one of the distinguishing features between Small Modular Reactors (SMRs) and traditional large-scale reactors in existing nuclear power plants. An SMR needs to possess load-following capability to integrate with non-dispatchable energy resources, e.g., renewable ones, to form an off-grid hybrid energy system. This paper provides a comprehensive review of different aspects of load-following operations of nuclear reactors by systemically examining early studies and some practical operating experience on existing nuclear reactors in various applications. These applications include electricity generation, marine vessel and spacecraft propulsion, and heat generation for industry applications. Considering different advanced reactor design concepts proposed for SMRs, the paper has concentrated on six representative reactor types and highlighted their unique features and feasibilities for load-following operations. Since an SMR can be considered as a combination of the reactor and the balance-of-the-plant, its power output can be regulated at the reactor power output or from the balance-of-the-plant. Different techniques to implement load-following operations have been described regarding different reactor designs and compositions of the balance-of-the-plant. Several special issues deserve special attention when performing load-following operations. These issues are rarely encountered in existing nuclear power plants for base-load operation, for example, imbalance in core power distribution, flow-induced vibration, excessive production, higher risk of fission products leaking, and flow-accelerated corrosion. The paper has provided necessary coverage of these less obvious consequences associated with load-following operations. • The load-following capabilities of an SMR is vital for off-grid power, marine, and space applications. • The choice of coolant and the reactor design can affect load-following characteristics. • Direct load-following techniques are constrained by physical limits of components. • Indirect load-following techniques can soften the stress on the reactor. • The load-following can exert stress on SMR components adversely. [ABSTRACT FROM AUTHOR]

Machine learning models are applied to simulated electrospray particle data to investigate plume dynamics from emission to final particle properties. A limited set of final particle properties are successfully regressed exclusively from emission property inputs. Random Forest model feature rankings for final plume angle reveal that particle charge has dominant influence when emission velocity is strictly axial, while lateral emission velocity has dominant influence when particles are emitted with an off-axis velocity component. In addition to providing correlations between initial and final particle properties, the machine learning models also identify correlations between different final particle properties. These correlations reveal opportunities for experimental approaches and diagnostic design by determining experimental measurements that offer insight into desired final particle properties. [Display omitted] • Machine learning enables rapid data interpretation for spacecraft propulsion. • Machine learning models assist electrospray diagnostic design optimization. • Machine learning models predict final particle properties from emission properties. • Random Forests strongly correlate charge with final angle in axial emission case. • Random Forests correlates off-axis lateral emission velocity with final angle. [ABSTRACT FROM AUTHOR]

Propellantless propulsion concepts based on electromagnetic waves like the EMDrive are claimed to be far superior with respect to the state of the art in spacecraft propulsion systems. Such devices consist of enclosed cavities with different geometric shapes that are injected with electromagnetic waves, producing unidirectional thrust without expelling propellant. Additional concepts emerged from theories like quantised inertia and involve laser-type EMDrives with optical cavity resonators and fiberoptic loops in the infrared spectrum. Claimed forces of these devices in the micronewton range are confronted with growing scepticism when basic conservation laws are applied. With cutting-edge measurement devices, we were able to characterize these concepts in a space-like environment with nanonewton resolution for thruster masses of up to 10 kg. Additionally, we enhanced our inverted double pendulum thrust balance with the ability to perform thrust measurements at cryogenic temperatures (65 K) to also operate a superconducting EMDrive that was claimed to have orders of magnitude higher thrust compared to classical resonators. In this paper, we present changes to each setup, based on criticism to our latest results, as well as thrust measurements of each device. Neither the EMDrive cavities nor the infrared laser resonators created a net-thrust above our balance noise. With the exception of the superconducting EMDrive, our data limits anomalous thrust below the threshold of classical propulsion with photon pressure for equivalent power-levels. Despite the enhancements made to each device, we did not detect any evidence in favour of the proposed theories. • Very precise thrust measurements of various EMDrive-Type of thrusters. • First implementation of the ability to test high temperature superconductors on sub-μN precision thrust balances. • Identification of false-positive effects and setting of limits for anomalous thrusts on these type of proposed concepts. [ABSTRACT FROM AUTHOR]

This paper investigates two strategies that exploit low-thrust propulsion and natural effects for the complete de-orbiting of spacecraft from a Low Earth Orbit (LEO). The first strategy aims to actively lower the perigee altitude by low-thrust propulsion to achieve passive drag-induced re-entry. The second strategy aims to actively move the spacecraft by low-thrust propulsion to reach a specific condition that can provoke passive orbital decay by means of the coupled effect of natural perturbations. For each strategy, a sub-time-optimal closed-loop steering law, which is proved to be stable, is designed with the Lyapunov method. Then a set of maps that show the costs of de-orbiting from LEO (i.e., the Δ v -budget and de-orbiting time) are plotted as a function of the initial conditions for the two strategies. In this way, the feasible initial conditions to apply the two strategies are identified by comparing the Δ v -budget. Before plotting the maps, the averaged low-thrust motion is derived, to reduce the computational load for the orbital propagation of low-thrust motion. • Investigating two de-orbiting strategies with low-thrust propulsion and natural perturbations. • Designing closed-loop and sub-optimal low-thrust steering laws. • Plotting de-orbiting maps as a function of initial conditions. • Identifying feasible initial conditions to apply the two de-orbiting strategies. [ABSTRACT FROM AUTHOR]

Gravitational Wave (GW) signal was first detected by Advanced Laser Interferometer Gravitational-Wave Observatory (aLIGO) in 2005 making GW detection a hotspot. The low-frequency gravitational signal needs to be detected in space. As the actuator of the drag-free control system, the micro propulsion system works to provide a super-stable platform for satellites by offsetting non-conservative forces. In this paper, some micro propulsion systems are analyzed from their work mechanisms and task requirements to justify suitability for space GWs detection. According to the analysis results, recent progress for cold gas propulsion, wave ionization ion propulsion, Hall propulsion and field emission effect electric propulsion (FEEEP) are introduced in detail. Lastly, the future of micro propulsion for space GWs detection is analyzed from engineering respects. The cusped magnetic field confines electrons by magnetic mirror effect. It means electrons would be exerted a force illustrated as equation (11), dragging them when they move to some regions with stronger magnetic field. Therefore, electrons are dragged when they move to the magnetic cusp and confined by this way. Most electrons can be confined between two magnetic cusps and the others can cross magnetic cusp by Bohm condition and near wall condition. The magnetic configuration can decrease the interaction between plasma and wall, prolong thruster lifetime and enhance discharge stability when the thruster is adjusted in large range [67]. • Analysis the requirements and the root of difficulties of micro propulsion system for space GWs detection. • Analysis the feasibility of micro propulsion systems for space GWs detection from their mechanism. • Sort out four micro propulsion systems potential for space GWs detection and introduce their recent progress. [ABSTRACT FROM AUTHOR]

The sloshing of liquids in low-gravity entails several technical challenges for spacecraft designers and operators. Those include the generation of significant attitude disturbances, the uncontrolled displacement of the center of mass of the vehicle or the production of gas bubbles, among others. Magnetic fields can be used to induce the reorientation of magnetically susceptible propellants and improve the controllability of a fluid system. Despite being proposed in the early 1960s, this approach remains largely unexplored. This paper provides new insight into the prospects and challenges of using magnetic control of space propellants. Key unanswered theoretical and technical questions are identified, highlighting the importance of developing appropriate analytical tools and fluid-magnetic simulation frameworks. New results associated with the reachability, scaling, long-term thermal and radiation stability, and efficiency of paramagnetic and ferromagnetic propellants are presented. Magnetic settling forces are shown to enhance the stability and speed up the oscillatory response of the liquid, leading to more predictable propellant management systems for different scales and filling ratios. These effects are particularly relevant for ferrofluids, whose enhanced magnetic properties make them excellent candidates for active sloshing control applications in space. • Magnetic forces can be used to control susceptible liquids in low-gravity. • Applications for space propulsion span from liquid reorientation to sloshing control. • The reachability, scaling, stability, and efficiency of this approach are addressed. • Key knowledge gaps are identified for paramagnetic and ferromagnetic liquids. [ABSTRACT FROM AUTHOR]

• This paper studies the space nuclear energy propulsion system using fission fragments for propulsion. • The uniform magnetic field model and the magnetic collimation structure model are proposed for analysis. • RMC and MATLAB were used to simulate the critical characteristics and propulsion performance of the system, respectively. • Using Am-242 m as fuel, the calculation finds that the system can achieve criticality with good safety. • The I SP of the magnetic collimation system can reach 6*105s, which is 40 % higher than that of the uniform model. Deep space exploration holds immense significance for the future of space exploration, and it requires propulsion systems with high specific impulse characteristics to accomplish long-duration space missions. This paper focuses on studying a space nuclear propulsion system based on the direct propulsion of fission fragments. Am-242 m is utilized as fuel, and two system models, namely uniform magnetic field and magnetic collimation system, are proposed. The system's critical and propulsion characteristics are analyzed using RMC and MATLAB, respectively. The calculation results demonstrate that the system with magnetic collimation exhibits superior comprehensive performance, with a specific impulse of approximately 6 × 105s. By adjusting the system materials and parameters, the k eff of the system is optimized to reach about 1.02, while the utilization rates of fission fragments and energy are approximately 20 % and 10 %, respectively. The results also indicate that the fission fragment direct propulsion system possesses excellent safety features and can play a crucial role in deep space exploration. Furthermore, the paper highlights the exploration of propulsion systems utilizing Pu239 and U235 in the future, as well as the investigation of enhancing the fission fragment escape rate and energy utilization rate. [ABSTRACT FROM AUTHOR]

• ZrC coatings have been fabrication at ORNL in a specialized coating system. • Coatings were characterized by SEM, Raman, nanoindentation, and XPS tools. • Coatings were uniformly ZrC crystals with some conditions creating C deposits. • Grains were primarily columnar with axis diameters from 0.3 – 2.3 µm. Coated fuel particle architectures with ZrC coatings are candidate fuels for advanced power reactors and space nuclear propulsion (SNP) concepts. Owing to its relevance to SNP, the composition, microstructure, and mechanical properties of eight ZrC coatings prepared by fluidized bed chemical vapor deposition were evaluated. Evaluation by SEM and EBSD showed that all grains were columnar. Across the various examined samples, minor axis diameters varied between 0.3 and 1.1 μm, and major axis diameters varied between 0.4 and 2.3 μm. Major and minor diameters increased with thickness particularly at higher deposition temperatures in which the major grain axis (from an ellipse fit to the grain shape) increased by 2.5 μm over the entire coating. Coatings with higher reactive gas flows and Zr/C concentrations closer to 1 were observed to contain nanocrystalline graphite deposits. Reactive gas flow doubling led to increases in coating thickness from around 10–15 μm to around 22–27 μm. [ABSTRACT FROM AUTHOR]

Phase-separation is of great importance in spacecraft tanks. The propulsion engines of spacecrafts require a single-phase propellant supply, which is a challenge using cryogenic liquids. In order to extend the knowledge of the capillary transport of cryogenic liquid in porous materials, we investigated wicking of liquid nitrogen at quasi-isothermal conditions in porous structures. The experiments were performed at saturation temperature in a one species system using the weight measurement technique. The setup configurations enabled to perform experiments with a sample overheat below 1 K. Porosity and permeability were calculated from the liquid nitrogen wicking experiments. The dimensionless representation of the cryogenic wicking experiments shows a good agreement compared to the wicking experiments with storable liquid. It leads to the statement that the combination of macroscopic parameters were obtained correctly for each experiment. • Wicking into porous material at ambient temperature with storable liquids. • Cryogenic wicking at quasi-isothermal conditions. • Dimensionless representation of wicking experiments. [ABSTRACT FROM AUTHOR]

Expanding the role of Space Physical Science and Technology in the modern world comes through the application of digital technologies and supercomputing. The present paper is aimed at getting the readers acquainted with results of resent scientific discussions on the topic having taken place at specialized scientific forums supported by the International Academy of Astronautics. The paper provides a breath coverage of results of main discussions and current trends in developing modern digital approach to solving acute problems of Space Physical Science and Technology. Numerical simulations with massive parallelization have natural limitation of efficiency. The results of recent studies demonstrate that each algorithm has an optimal number of parallel threads providing maximal efficiency. On further increasing the number of parallel processes, one encounters a decrease of efficiency. The universal supercomputers can hardly provide high efficiency in solving multi-scale problems. Creating supercomputers of a specialized architecture focused on a narrow class of multiscale problems will allow for efficient parallel solution of such problems. • Protection of Space structures from space debris collisions and micrometeoroids. • Fire safety of Space vehicles and Propulsion systems. • Supercomputer predictive modeling for ensuring Space program safety. • Each algorithm has an optimal number of parallel threads providing maximal efficiency. [ABSTRACT FROM AUTHOR]