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Electric propulsion (EP) has become one of the most promising options for the motion control of small satellites. The physics behind the ejected plasma plumes of EP thrusters has attracted significant interest due to their interactions with the critical components of satellites. Axisymmetric plume assumptions are widely used in simulations and plume diagnostics. However, we show here that the plasma plume of a parallel-plate pulsed plasma thruster (PPT) is asymmetrically distributed along the centerline of the electrodes, contrary to the inherent axisymmetric assumption. To study this asymmetric plume structure in depth, a triple Langmuir probe was used to obtain the electron density of a two-dimensional plume area over the operating period of a PPT. The electron density results show that the plasma forms an 'I' shape plume at 2 μs after the initial main discharge. However, over time, the plume appears to cant significantly towards the cathode. The physical mechanism behind the asymmetric plume structure is studied through inter-electrode magnetic probe measurements and plasma trajectory analyses. The successive magnetic profiles indicate that the plasma is accelerated by a non-symmetrical electromagnetic force between the electrodes, which results in the plasma exhausting out with an entirely asymmetric distribution shifted upwards towards the cathode. This was also verified using varying sets of discharge voltage experiments. This work also indicates that care must be taken in the selection of the measurement points in PPT plume diagnostics. The measurement points should be chosen above the centerline of the PPT exit as the plasma parameters along the centerline may not be the most energetic part as previously believed. Furthermore, the asymmetric acceleration component of the electromagnetic force between the electrodes can enlighten us in the design of electromagnetic field configurations for future discharge channel optimization. [ABSTRACT FROM AUTHOR]

The ablative pulsed plasma thruster (APPT) is a typical form of electric propulsion that is highly suitable for micro-satellites. However, its low performance has been a main drawback. Recently, an asymmetric segmented anode schematic was shown to enhance the performance of APPTs. The improved thrust performance obtained with the use of a segmented anode can potentially enable promising applications in future space tasks. To further understand the physical processes behind a segmented anode APPT, experiments using an ultra-high-speed camera, narrow bandpass filters, and a magnetic probe were conducted with both normal parallel-plate electrodes and a segmented anode on an experimental APPT. Successive images of light emission from C+, neutral C2, and broadband emission reveal the evolution of the plasma morphology on an APPT with a segmented anode. The magnetic field strength profiles show that the inter-electrode plasma propagation velocity of a segmented anode APPT is approximately 19 km s−1, which is slightly lower than that of a normal APPT at 23 km s−1. However, the similar order of magnitude suggests a similar downstream acceleration mechanism between the two. The arc morphology observed in the high-speed images shows that the segmented anode APPT with asymmetric electrodes has a restricted arc attachment point, possibly resulting in stronger arc current density at the anode corner. This should result in higher ionization and ablation when compared with a normal anode APPT. Based on the results here, we hypothesize that the segmented anode APPT may be able to better utilize the restrike energy in the propellant ionization process. This fundamentally changes the physical process behind APPT operation, meaning that the improvements should also be applicable to other APPT designs. It can also be adopted as an additional step in the overall optimization of an APPT design. [ABSTRACT FROM AUTHOR]

In this work, a novel method of directly measuring the roll torque of low power Hall-effect thrusters is proposed. The measurement device uses an elastic pendulum with the thruster mounted on the same axis as the rotation axis of the pendulum, and the torque of the thruster is reflected in the rotation of the movement part and amplified by the pendulum arm with the measuring range of approximately 0–200 μN m with a resolution of less than 0.5 μN m. The roll torque measurement results of a Hall-effect thruster show that the rotational direction is consistent with the ion deflection direction by the magnetic field, and the roll torque ranges about 16–23 μN m at the anode power range of 300–450 W with cathode keeper, and the values is influenced by the cathode keeper currents. An empirical formula for the Hall-effect thrusters is proposed according to the reliable measurement results of the low power Hall-effect thruster, and the roll torque values are estimated for several low power to medium power Hall-effect thrusters. The method also provides a new idea to accurately evaluate the roll torque of other types of electric propulsion thrusters. • Firstly proposed a direct measurement method of roll torque for Hall-effect thrusters and other EP thrusters. • The first direct measurement to obtain the azimuthal force of a low power Hall-effect thruster. • Firstly proposed the roll torque estimation method of Hall-effect thrusters based on actual measurement results. [ABSTRACT FROM AUTHOR]

The Pulsed Plasma Thruster (PPT) was the first electric propulsion device to be implemented on a spacecraft in orbit (Zond 2) in 1964. However, the pulse ignition mode and the shot-to-shot discharge differences introduce challenges into the evaluation of the performance of PPTs. This work performs impact pressure measurements on a coaxial PPT (Pulsed Electric ThRuster of the University of Stuttgart, PETRUS) using a pressure probe. It is aimed to examine two specific thruster performance values that conventional thrust stands cannot easily obtain: the transient behavior during the phase where the final propellant surface contour is established (i.e., the shot-to-shot difference for the impulse bit during this phase) and the initial operation stability for the lifetime test. The impact pressure difference was measured at various discharge voltages using the pressure probe. To test the thruster's operational stability, a 5000 shots burn-in experiment was conducted on PETRUS. During the operation, the impulse bit variations were measured over 5000 continuous shots using the pressure probe. From the results, the impulse bit curve of PETRUS first shows a trend of decreasing impulse bit (the initial several hundred shots) and then an increasing trend (after 300 shots). After 2151 shots, the thruster's operating condition reaches a steady state with an impulse bit of 144 μN s. Additionally, the reliability of the pressure probe measurement is also evaluated by comparing the results with those from a thrust stand. The results are in good agreement at higher discharge voltages (from 0.9 kV to 1.6 kV), the maximum measured impulse bit difference of the pressure probe over the thrust stand is 9.7%. These results provide a dynamic evaluation of the thruster's operational stability from an engineering aspect. Correspondingly, the comparison results between the pressure probe and thrust stand is a preliminary verification of the pressure probe impulse bit measurement method in this work. • This work proposed a impact pressure measurement on coaxial PPT with a pressure probe. • 5000 shots burn-in experiment was conducted to test the thruster's operational stability. • This probe can measure the shot to shot dynamic pressure and impulse at short time scales (∼1s). [ABSTRACT FROM AUTHOR]

Increasing interest in development of very low Earth orbit (VLEO) has attracted more and more researchers to study atmosphere-breathing electric propulsion (ABEP) system in past several decades. This system can use rarefied atmospheric particles as the propellant of electric thrusters, and maintain a long lifetime mission without carrying any propellant from ground. As the key component of system, intake device can realize the collection and compression of atmospheric particles within limited frontal area, which determines the performance of whole ABEP system. This review summarizes the previous studies to develop intake devices, evaluates the corresponding performance and understands the model involved, including atmosphere model, flow physic model and so on. In addition, several continued researches for intake device are also presented, including ground experiment technologies, intake surface material development, space compressor and liquefaction technology. Wherever possible, comments have been provided to provide useful reference to researchers engaged in intake device for ABEP system. [ABSTRACT FROM AUTHOR]

Recently, an increasing number of small satellites have chosen electric propulsion (EP) as the propulsion system for motion control in space flights. Research on the ejected high energy plasma plumes from EP thrusters has attracted significant interest due to their interactions with the components of satellites. This work demonstrates a probe array diagnostic platform (with 67 Faraday probes on the array) that enables the acquisition of the full three-dimensional morphology of the plume current density from the near-field to the far-field of EP stationary thrusters. The plume parameters of an ion thruster including the 3-D plume morphology, divergence angle, and thrust vector angle are measured using this diagnostic platform. From 250 to 700 mm from the thruster exit, the plume field exhibits a stable divergence angle of 43.6°. By fitting circles to the contours of each cut plane, the average thrust vector angle is calculated to be 1.3°, and the repeatability error for the circle center fitting is approximately 2.2%. The current density profiles on exit cut planes demonstrate an "elliptical distortion" phenomenon in the plume shape. Using the ellipse fitting method, it is found that an elliptical shape was formed in the near-field of the plume. As the plume propagates to the far-field, the distortion of the ellipse becomes more apparent (the aspect ratio is 1:1.14). Some possible causes may be the geo-magnetic field, background vacuum environment distribution, or thruster manufacturing error. The thrust vector data can provide a basis for the control strategy of the satellites. In addition, the 3-D plume current density data can give us a comprehensive understanding of the plume plasma physics. These results may also provide an accurate and rich database for parameters for future plume simulations. [ABSTRACT FROM AUTHOR]