2D hybrid-PIC simulation of the two and three-grid system of ion thruster

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Beam flatness is an important parameter that determines the performance and the lifetime of a gridded ion thruster. To improve the beam flatness of the 30 cm (LIPS-300) ion thruster, variable aperture ion optics that adapts to the decreasing ion density as the radius increases is proposed. It is the ion optics that the screen grid surface is divided into several zones, where the aperture diameter in each zone is determined by the ion density and the electron temperature upstream of the screen grid. The beam current density in the central area is artificially reduced. A particle in cell-Monte Carlo collision model is applied in this work to investigating the effect of variable aperture on the perveance and the maximum beam current per aperture by simulating the extraction, focusing and acceleration processes of ions. Taking into account the engineering implementability, the screen grid surface is divided into four zones. The hole diameter in each zone is decreased from 1.95 mm to 1.8 mm, 1.9 mm, 1.8 mm and 1.7 mm, respectively. The simulation results show that the maximum ion density in the center area of grid is decreased by 10.6% and 6.99%, while it is increased by 6.49% and 22.3% in the edge region, respectively. The beam flatness of the variable aperture ion optics is improved from 0.69 to 0.88. The erosion rate is decreased by 31.9%, but the total beam current is also decreased by 7.15%. The simulation results can provide a valuable reference of the development of the ion thruster.

Section: Introductionmentioning

confidence: 99%

Investigation of variable aperture on the performance and lifetime of ion thruster

Chen

Jia

Geng

et al. 2021

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“…The magnetic field is simulated with and without considering the bending part. As shown in figure 7, the magnetic flux density is higher when the magnetic field induced by the bending structure is considered, compared to the one without the bending part, especially at the upstream of the discharge channel ranging from x=0 m to x=0.02 m. It could be seen from the successive image of C + in the APPT firing [12] that the ionization process happens near the surface area of the propellant, then plasma is accelerated from the surface. So the increase of magnetic field at the beginning of the discharge channel could enhance the acceleration of the charged particles near the propellant.…”

Section: Magnetic Profilementioning

confidence: 94%

“…The magnetic field might also be responsible for the erosion phenomenon due to its confinement of the ions that bombard the electrodes. When it comes to numerical study, an accurate magnetic field model could be the foundation of the magnetohydrodynamics or particle-in-cell simulations [11,12]. This work builds a model to describe the magnetic field in spatial and temporal view for a better understanding of the working process and thruster structural design.…”

Section: Introductionmentioning

confidence: 99%

Investigation of the self-induced magnetic field characteristics in a pulsed plasma thruster with flared electrodes

Liu¹,

Yang²,

Huang³

et al. 2019

The self-induced magnetic field in a pulsed plasma thruster (PPT) with flared electrodes is investigated for a better understanding of the working process and the structural design of the thruster. A two-dimensional model of the magnetic field is built and is validated by comparing the simulated results with the experimental results in literature. The magnetic flux density in the discharge channel during the working process is presented and analyzed regarding the electrode structures. The calculated magnetic field flux density decreases from 0.8 T at the upstream to 0.1 T and below at the downstream in the discharge channel (68 J). The peak of the magnetic flux density over time lags behind the current peak, which provides evidence for the existence of a moving plasma sheet in the discharge process. The magnetic field induced by the current in the extra bending part of the anode enhances the Lorentz force, which acts on the charged particles near the propellant. Finally, the geometric study indicates that the electromagnetic impulse bit does not monotonically increase with the flared angle of the electrodes. Instead, it reaches a maximum at a certain flared angle, which could provide significant suggestions for structural optimization.

“…Ion thrusters [1][2][3] have the potential to dramatically enhance the payload, prolong the service life and reduce the launching budget of spacecraft owing to their high specific impulse, low thrust, prolonged lifespan, and repeatable start-up etc, and have been successfully employed in missions concerning attitude control, station keeping, drag compensation, and deep space exploration. The implementation of electric propulsion technology on spacecraft will alter the plasma environment in the vicinity of the spacecraft owing to its special plume environment, resulting in a series of effects such as charged spacecraft surfaces, electromagnetic interference, and subsequent changes in the thermal properties of thermally sensitive surfaces.…”

Section: Introductionmentioning

confidence: 99%

Investigation into the thermal effect of the LIPS-200 ion thruster plume

Chen¹,

He²,

Zuo³

et al. 2022

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The distributions of the thermal effects of the ion thruster plume are essential for estimating the influence of the thruster plume, improving the layout of the spacecraft, and the thermal shielding of critical sensitive components. In order to obtain the heat flow distribution in the plume of the LIPS-200 xenon ion thruster, an experimental study of the thermal effects of the plume was conducted in this work with a total heat flow sensor and a radiant heat flow sensor over an axial distance of 0.5–0.9 m and an angle of 0°–60° of the thruster. Combined with a Faraday probe and a retarding potential analyzer, the thermal accommodation coefficient of the sensor surface in the plume is available. The results of the experiment show that the xenon ion thruster plume heat flow is mainly concentrated within the range of 15°. The total and radial heat flow of the plume downstream of the thruster gradually decreases along the axial and radial directions, with the corresponding values of 11.78 kW/m2 and 0.3 kW/m2 for the axial 0.5 m position respectively. At the same position, the radiation heat flow accounts for a very small part of the total heat flow, approximately 3%–5%. The thermal accommodation factor is 0.72–0.99 over the measured region. Furthermore, the PIC and DSMC methods based on the Maxwell thermal accommodation coefficient model (EX-PWS) show a maximum error of 28.6% between simulation and experiment for LIPS-200 ion thruster plume heat flow, which on one hand provides an experimental basis for studying the interaction between the ion thruster and the spacecraft, on the other hand provides optimization of the ion thruster plume simulation model.