Journal of electric propulsion最新文献

The performance of an innovative dielectric bimaterial is assessed for use in Hall-effect thrusters. The bimaterial consists of a structural body of graphite with a converted surface layer of hexagonal boron nitride (h-BN). The bimaterial couples the dielectric behavior and low emissivity of h-BN at its surface, with the thermal shock resistance and machinability of graphite at its core. In this paper, the performance of graphite/h-BN bimaterials synthesized from liquid-phase and vapor-phase carbothermic reactions of B₂O₃ in nitrogen is compared and evaluated against the state-of-the-art wall material, bulk h-BN. Graphite/h-BN bimaterials synthesized through vapor-phase carbothermic reactions are shown to perform comparatively better than bimaterials synthesized through liquid-phase carbothermic reactions. The erosion rate of vapor-phase grown h-BN layers is similar to that of bulk h-BN with a sputtering yield of 0.021 mm³/C for xenon ions at 300 V.

Supplementary information: The online version contains supplementary material available at 10.1007/s44205-025-00126-0.

Electric propulsion systems require careful consideration of plume divergence and evolution over a range of operating conditions and environments. Existing means of describing plume divergence such as outlines, plume profiles, and snapshots of the plume are dominated by outlier particles and do not provide reliable or quantitative insight. Proposed herein are two novel methods for describing plume divergence using standard deviation and emittance to provide quantitative insight of the collective behavior of plume species. Furthermore, the emittance metric from the particle accelerator community is shown to accurately describe plume evolution in a two-dimensional position and momentum angle space. Cross-sectional emittance measurements are used to display the presence of non-Hamiltonian forces in plume evolution, namely stochastic Coulomb collisions between neighboring particles. Finally, full-plume emittance diagrams are demonstrated as a means of identifying when an electric propulsion plume has reached steady state.

Magnetic shielding technology has extended the operating life of Hall thrusters to timescales enabling them to expand beyond missions in Earth orbit, which typically require less than 10 kh of operation, to deep-space missions requiring 10-30 kh of operation. While this has expanded potential mission capture, the specific impulse of flight Hall thrusters is still less than 2,000 s, which can limit their use on trajectories requiring velocity change greater than 10 km/s. To expand mission capture further, specific impulses greater than 3,000 s are needed over wide power throttling ratios. Towards those ends, JPL has developed a low-mass, 10-kW class, 3,000 s specific impulse Hall thruster that has demonstrated a 2:1 power throttling ratio at 800 V (~ 3,000 s specific impulse), efficiencies greater than 50% over 6:1 power throttling, and a 50:1 power throttling ratio over 0.2-10 kW. Thermal steady-state operation was reached at 15 kW, and a maximum operating power of 20 kW was briefly demonstrated. To achieve this performance requires novel approaches that can operate at power densities significantly exceeding the state-of-the-art. High power density operation is achieved in the H10 Hall thruster with an integrated, conducting wall, magnetically shielded discharge chamber assembly combined with a passive, multi-zone heat rejection system. Testing has demonstrated peak performance at 800 V, 10 kW of 457 mN thrust, 3400 s specific impulse, and an efficiency of 76%. The H10 represents a new class of high-power density Hall thrusters, enabling next generation robotic science and human exploration missions.

In the field of low power (< 100W) electric propulsion, all thruster options demand significant trade-offs between operating parameters, systemic simplicity and scalable cost. With many systems on the market there are concerns with reliability, the need for extra considerations such as electron neutralizers for pure ion plumes (the flight proven PPTs and this Metal Plasma Thruster (MPT) generate quasi-neutral plasmas and are exceptions), and the cost or craft compatibility of propellants. The Metal Plasma Thruster (MPT) is a new type of electric propulsion technology intended for low power applications. The system imparts momentum using inert, solid metal pucks as a propellant by using pulsed power to convert the metal into high velocity (~ 17 km/s for Mo) jets of quasi-neutral plasma. The MPT technology does not require gas or liquid propellants, neutralizers, standby heaters, high voltage electronics, high electric or magnetic fields to operate. This comparatively simple pulsed operation is amenable to closed loop control, which provides for fine thrust control and S/C directed impulse on demand. Furthermore, the technology can use any metal as propellant, opening up unique opportunities for In-Situ Resource Utilization (ISRU) as well as customizability of performance for meeting specific mission needs. This paper describes implementation and direct measurement of this closed loop control mode as well as impulse measurement of multiple metals consistent with the aim of ISRU at NASA Glenn Research Center (GRC).

The feasibility of a novel 'magnetic arch' topology for controlled contactless plasma beam acceleration is experimentally demonstrated using a pair of coaxial electron cyclotron resonance sources with opposing magnetic polarities, such that their respective magnetic nozzles connect to form a closed-line configuration. Retarding potential analyser measurements are taken for a single source and the two sources with the same and opposing polarities, as well as no applied magnetic field, showing that the magnetic arch yields higher maximum ion current and lower plume divergence angle than other alternative configurations, albeit the most probable ion energy is lower than for a single magnetic nozzle, in agreement with existing models. This validation paves the way to clustering magnetic-nozzle-based plasma thrusters for space propulsion.

Electrospray thruster ground testing, with well understood facility effects, is of critical importance to qualify the technology for long duration flight missions. While there has been substantial work to understand the beam physics and plume dynamics of electrospray thrusters and the implications thereof on performance and lifetime, work to understand the impact of facility effects has been neglected until recently. Interactions between an electrospray plume and the vacuum chamber test facility have implications on both performance and lifetime. Therefore, any effort to characterize electrospray thruster performance and lifetime must be done so with an understanding of facility effects. In some ways, this is no different than the significant investment that has been made to understand the facility effects for plasma thruster testing. However, there are different challenges with the management of positively charged, negatively charged, and neutral propellant particles across a distribution of particle charge and mass when testing electrospray thrusters in a vacuum chamber. The focus of this paper is to characterize the significance of secondary particles from the impact of ionic liquid electrosprays with a beam target, and the influence of a novel beam target design and biasing. Results on secondary current and mass flux measurements are presented with some initial results on secondary time-of-flight measurements from the beam target. Additionally, beam target modeling results are presented to support the experiments and interpretation of the results. The results revealed secondary particles with an average charge-to-mass ratio as low as 31 C/kg, and that an improperly biased beam target, or no beam target, can artificially inflate emitted current due to electron back streaming by as much as 20%. The experimental and modeling results suggest an optimized beam target and screen voltage of -100 V and -200 V, respectively. If no consideration of facility effects is included in testing electrospray thrusters, performance, reliability, and lifetime can be adversely affected, and premature thruster failure may result. The work presented here improves our understanding of facility effects and our capabilities to mitigate them to successfully qualify and acceptance test electrospray thrusters for flight.

A porous conical type electrospray array thruster consisting of 6102 individual emitters is operated at up to 13.3 W power. The design and manufacture of the thruster are described, including its porous glass emitter chip and metallized ceramic extractor chip. A precision mass balance mounted inside a bell jar is used to directly measure the thrust, specific impulse, and efficiency in negative polarity, from $42\pm 0.5\mu$ N, $1050\pm 26$ s, and $57\pm 1.9$ % at $-1000$ V and 0.38 W to $174\pm 0.5\mu$ N, $420\pm 2$ s, and $21\pm 0.3$ % at $-1300$ V and 1.7 W. Additional negative polarity experiments in a 2 meter vacuum facility demonstrate powers from order 1 $\mu$ W to over 10 W, spanning 7 orders of magnitude. Power and performance measurements were not repeated for positive mode operation, as this was found to induce arcing between the emitter and extractor electrodes at 1400 V and above. The drop in efficiency from $-1000$ V to $-1300$ V operation in the bell jar is discussed within the context of facility effects, and secondary charged particle flux to the thruster is identified as a likely contributor. Finally, the performance of the thruster is considered relative to scaling electrospray systems to higher power robustly.

Supplementary information: The online version contains supplementary material available at 10.1007/s44205-025-00114-4.

The collisionless cross-field electron transport in an $E\times B$ plasma configuration, representative of a Hall thruster, is studied using bispectral analysis on the data of a fully-kinetic simulation. The nonlinear, in-phase interaction of the oscillations of the azimuthal electric field and the electron density, both tied to the fundamental electron cyclotron drift instability (ECDI) mode, is found to be the main driver of electron transport. Higher-wavenumber ECDI modes do not drive anomalous transport directly; however, they are nonlinearly coupled with each other and with the fundamental ECDI mode. In addition, there is a smaller contribution from a lower-wavenumber mode, not predicted by linear ECDI theory. A reduced model obtained by sparse regression of the data suggests the existence of an inverse energy cascade from the higher ECDI modes to the fundamental one, which would mean that these modes do contribute to transport, albeit indirectly.