Coaxial Pulsed Plasma Thruster Research Articles

Influences of initial voltage and electrode size on propellant surface evolution in a coaxial pulsed plasma thruster

Coaxial pulsed plasma thrusters are well suited for microsatellites, but the discharge therein is radially nonuniform, resulting in uneven propellant ablation that ultimately degrades the thruster performance. This study investigates how the propellant surface evolves during 50 000 discharges in a coaxial pulsed plasma thruster with different outer electrode inner diameters and initial voltages, encompassing the propellant surface profile, concave depth of the propellant, and carbon deposition on the propellant surface. The results show that the propellant surface changes in real time throughout the 50 000 discharges. Increasing the outer electrode inner diameter leads to a proportional increase in the concave depth of the propellant, resulting in unstable ablation, and the increased inward concave depth increases the likelihood of discharge failure. Furthermore, increasing the initial voltage increases the energy density in a unit area of propellant, further contributing to the increase in the concave depth and discharge failure. The energy density distribution on the propellant surface is obtained via simulation, and by comparing the results with experimental data, the critical ablation threshold for the propellant is estimated as being ca. 0.7 J/mm2.

Study on the influence of ignitor position on a coaxial pulsed plasma thruster

Because of their unique advantages, coaxial pulsed plasma thrusters are competitive propulsion systems for microsatellites. The ignitor of a coaxial pulsed plasma thruster can be mounted on the side or at the center, but previous studies have not compared the effects of the two mounting positions on the thruster. Therefore, this study designs and investigates a coaxial pulsed plasma thruster that permits the ignitor position to be changed while keeping the electrode geometry constant. The discharge circuit parameters, plume plasma characteristics, and propulsion performance are measured, and long-exposure photographs are taken. While the impulse bit of the thruster with a coaxially-mounted ignitor is slightly lower than with a side-mounted ignitor (319.5 μNs vs. 334.5 μNs), the ablated mass bit of the former is significantly lower (15.2 μg vs. 24 μg). The thruster with a coaxially-mounted ignitor achieves higher specific impulse and efficiency (2139 s and 16.74%). Because of the match between the electric field and the neutral gas supply during discharge in the thruster with a coaxially-mounted ignitor, its energy and propellant utilization is higher, thus achieving a better propulsion performance.

A Coaxial Pulsed Plasma Thruster Model with Efficient Flyback Converter Approaches for Small Satellites

Pulsed plasma thrusters (PPT) have demonstrated enormous potential since the 1960s. One major shortcoming is their low thrust efficiency, typically <30%. Most of these losses are due to joule heating, while some can be attributed to poor efficiency of the power processing units (PPUs). We model PPTs to improve their efficiency, by exploring the use of power electronic topologies to enhance the power conversion efficiency from the DC source to the thruster head. Different control approaches are considered, starting off with the basic approach of a fixed frequency flyback converter. Then, the more advanced critical conduction mode (CrCM) flyback, as well as other optimized solutions using commercial off-the-shelf (COTS) components, are presented. Variations of these flyback converters are studied under different control regimes, such as zero voltage switching (ZVS), valley voltage switching (VVS), and hard switched, to enhance the performance and efficiency of the PPU. We compare the max voltage, charge time, and the overall power conversion efficiency for different operating regimes. Our analytical results show that a more dynamic control regime can result in fewer losses and enhanced performance, offering an improved power conversion efficiency for PPUs used with PPTs. An efficiency of 86% was achieved using the variable frequency approach. This work has narrowed the possible PPU options through analytical analysis and has therefore identified a strategic approach for future investigations. In addition, a new low-power coaxial micro-thruster model using equivalent circuit model elements is developed.This is referred to as the Carlow–Stuttgart model and has been validated against experimental data from vacuum chamber tests in Stuttgart’s Pulsed Plasma Laboratory. This work serves as a valuable precursor towards the implementation of highly optimized PPU designs for efficient PPT thrusters for the next PETRUS (pulsed electrothermal thruster for the University of Stuttgart) missions.

Study of a low-power micro coaxial pulsed plasma thruster with axial applied magnetic field

Pulsed plasma thrusters (PPTs) have been widely used in nano/micro-satellites in recent years. However, the efficiency of a low-power (<10 W) PPT is typically lower than 10%. Applied magnetic fields are widely used in electric propulsion to improve performance. However, few attempts have been made to use applied magnetic fields in low-power PPTs because of the high discharge current (kA), short pulse duration (μs), and oscillating current in the discharge process. The results of an experimental study on a low-power micro coaxial PPT with axial applied magnetic field are presented in this paper. The thrust, ablation, and discharge characteristics of the thruster are obtained under different applied magnetic fields. The results show a significant improvement in thruster efficiency. Using magnets that can generate magnetic fields of up to 0.3 T, 0.4 T, and 0.5 T result in efficiencies for the PPT of 15.6%, 24.7%, and 67.8%, respectively, which are 1.91, 3.03, and 8.34 times that in the absence of an applied magnetic field (8.1%). The results also show that the applied magnetic field has little influence on the discharge circuit and slightly affects the discharge characteristics and ablated mass bit (the consumption of propellant per discharge) of the thruster.

Propellant Surface Evolution of a Coaxial Pulsed Plasma Thruster

Coaxial pulsed plasma thrusters are suitable for small satellites. In such thrusters, the discharge is radially nonuniform, which leads to an ablation nonuniformity that causes variations in the propellant surface. This paper describes a study of the propellant profile and ablated mass bit (i.e., mass ablated from the propellant in each discharge) of a coaxial pulsed plasma thruster over 100,000 discharges. The results show that the ablated mass bit of the thruster increases initially before becoming confine within a certain range. The ablated mass bit is less than over the first 2000 discharges, and then stabilizes at when the number of discharges exceeds 5000. In addition, the propellant profile changes over the 100,000 discharges, and the maximal concave depth remains less than 3 mm. Visual observations show that the area of carbon deposition is relatively large at a low discharge number. The propellant ablation area is an important factor in the ablated mass bit, because its variation is similar to that of the ablated mass bit during the continuous discharge process.

Investigation on operational stability of a pulsed plasma thruster with a pressure probe

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.

Influence of Anode Diameter for Lifetime of Coaxial Pulsed Plasma Thruster

Influence of Anode Diameter for Lifetime of Coaxial Pulsed Plasma Thruster

Impulse bit and ablation characteristics of double cylindrical pulsed plasma thruster

Previous studies have shown that the performance of coaxial pulsed plasma thrusters depends on the ratio of propellant surface area-to-cavity volume. In this work, a double cylindrical propellant was introduced to increase the operation ratio compared to coaxial propellants; the results demonstrated that the geometric parameter obtained by dividing the ratio by the hollow diameter greatly influences the impulse bit. Also, direct observations and mass shot measurements revealed a higher density of the ablation gas from the cylindrical part than from the hollow part.

Study and modeling of propellant ablation in coaxial ablative pulsed plasma thrusters

The influence of the main design parameters on the ablated mass bit (the mass ablated from the propellant in each discharge) of a coaxial ablative pulsed plasma thruster is analyzed in this paper. The analysis results show that the main influencing factor of the ablated mass bit in coaxial ablative pulsed plasma thrusters is the ablation energy from the arc. When the ablation energy is given, the shape and size of the propellant surface mainly affect the distribution of the ablated mass but have less influence on the ablated mass bit. Based on the analysis, a simple model that can be used to estimate the ablated mass bit of the coaxial ablative pulsed plasma thruster is established. To evaluate and complete the new model, a coaxial ablative pulsed plasma thruster is operated and tested under 10 different working conditions. The experimental results show that the ratio of the ablated mass bit to the absolute value integrals of the product of the electrode voltage and discharge current is approximately 0.38–0.49 μg/J. When the ratio is set to 0.44 μg/J, the maximum and average errors of the new model are 15.7% and 7.6%, respectively, for ablated mass bits under these 10 working conditions. The concept of the ablation direction is proposed in this paper. The experimental and simulation results show that the ablation direction has a great influence on the propellant surface profile change in coaxial ablative pulsed plasma thrusters, which also affects their performance and reliability.

A model for macro-performances applied to low power coaxial pulsed plasma thrusters

Pulsed plasma thrusters (PPT), as an important developing direction of micro-satellite propulsion system, are attracting more and more attention. Low power coaxial PPT has great advantages in miniaturization because of its coaxial structure, which is beneficial for micro-satellite applications. Since there are few macro-performances estimation model in low power coaxial PPT, this paper establishes a model based on one-dimensional electromechanical PPT model and gas dynamic pressure forces model to describe macro-performances of low power coaxial PPT. In this model, the electromechanical model describes the plasma electromagnetic acceleration process, current and voltage changes of the main circuit during the discharge process, and the gas dynamic pressure forces model describes the expansion and acceleration of the neutral gas group in the discharge channel. To verify the model, we compared the voltage and current obtained from model simulation with the experimental results, and compared the impulse bit measured in the experiment with the impulse bit calculated by the model. The results show that the model established in this paper can well simulate the variation of voltage and current during the working process of low power coaxial PPT, and can preliminarily estimate the macro performances of the thruster.

Firing test and evaluation of coaxial pulsed plasma thruster

For many reasons, electrical propulsion systems are the most expected in space; they are suitable for researchers because they are effective, easy to design and applicable in space. One of the most important parameters that could assess the performance of the propulsion system is the problem of measuring the very low thrust of pulsed plasma thrusters (PPTs). Various test beds, including indirect designs of measuring systems such as pendulums and optical measuring systems, can measure the thrust with very high precision, but it is very complicated.This paper presents the firing test of a new designed coaxial pulsed plasma thruster for proofing the concept and then the thrust is measured by a new simple test bed using a rod in accordance with a loading cell (0-100 gf.) which can measure thrust accurately after good calibration up to μgf. In a medium-sized vacuum chamber (VC) (10−5 bar, D= 1 m and L= 1.4 m), a coaxial PPT is designed and tested under environmental conditions in space. Four proofs of concept took place until the experiment was adjusted and six experiments were carried out to measure the thrust using the new test bed under various operating conditions. Evaluation of the system is presented too and the maximum thrust obtained is 0.68mN at the discharge voltage equals 7kV and the 12kV voltage of the spark plug initiator and the efficiency equals 10%. The steps of fabrication of the coaxial PPT and steps of measuring the thrust is presented in the paper beside the result charts of the thrust under different operating conditions.

Study on the plasma generation characteristics of an induction-triggered coaxial pulsed plasma thruster

At present, spark plugs are used to trigger discharge in pulsed plasma thrusters (PPT), which are known to be life-limiting components due to plasma corrosion and carbon deposition. A strong electric field could be formed in a cathode triple junction (CTJ) to achieve a trigger function under vacuum conditions. We propose an induction-triggered electrode structure on the basis of the CTJ trigger principle. The induction-triggered electrode structure could increase the electric field strength of the CTJ without changing the voltage between electrodes, contributing to a reduction in the electrode breakdown voltage. Additionally, it can maintain the plasma generation effect when the breakdown voltage is reduced in the discharge experiments. The induction-triggered electrode structure could ensure an effective trigger when the ablation distance of Teflon increases, and the magnetic field produced by the discharge current could further improve the plasma density and propagation velocity. The induction-triggered coaxial PPT we propose has a simplified trigger structure, and it is an effective attempt to optimize the micro-satellite thruster.

Numerical simulation of the coaxial magneto-plasma accelerator and non-axisymmetric radio frequency discharge

The paper presents the results of mathematical modeling of physical processes in electronic devices such as helicon discharge and coaxial pulsed plasma thruster. A mathematical model of coaxial magneto-plasma accelerator (with a preionization helicon discharge), which allows estimating the transformation of one form of energy to another, as well as to evaluate the level of the contribution of different types of energy, the increase in mass of the accelerated plasmoid in the process of changing the speed. Main plasma parameters with experimental data were compared.

Plume Characterization of Electric Solid Propellant Pulsed Microthrusters

Electric solid propellants are an attractive option for propulsion because they are ignited only by an applied electric current. This leads to capabilities such as on-demand throttling and reignition in chemical combustion operation, and it lends itself to application in electric propulsion. In this work, the plasma plume created by electric solid propellant pulsed microthrusters is investigated using a nude Faraday probe, a triple Langmuir probe, and residual gas analysis. The thrusters are tested at a vacuum level of . Results indicate an average plume electron temperature of 1.7 eV and a density of to . A time-of-flight analysis indicates an exhaust velocity of on the centerline. Additionally, ionization fraction is estimated to be 0.3% for an ablation mass bit of on average. Thermal equilibrium and adiabatic expansion calculations suggest a nonequilibrium gas with high-temperature electrons and lower-temperature neutral and ion species. Results indicate the pulsed microthruster operates similar to a coaxial pulsed plasma thruster, which is dominated by electrothermal effects.

Study on the Discharge Characteristics of a Coaxial Pulsed Plasma Thruster

A new coaxial pulsed plasma thruster (PPT) laboratory model is designed and employed in this study. A Teflon sleeve is connected with the anode, which is shaped as a nozzle, and a cathode is mounted in the cavity of the Teflon sleeve and kept in close contact with it. A thread is then designed in the internal surface of the Teflon sleeve, and because of the strong field strength of the cathode triple junction (CTJ), vacuum flashover occurs and a plasma jet is acquired behind the anode. The electric field distribution of the designed coaxial PPT laboratory model is simulated by MAXWELL 3D simulation software, and the plasma density and thrust are measured by a Langmuir probe and a piezoelectric thin-film sensor, respectively. Through a series of comparative experiments, we discuss the impact of optimal designs, such as the thread and the nozzle-shaped anode, on the discharge characteristics of the coaxial PPT. The experimental and simulation results indicate that the designed coaxial PPT laboratory model presents better discharge characteristics in view of its higher plasma density and greater thrust.

Effect of Capacitance on the Performance of a Di-methyl Ether Pulsed Plasma Thruster

We have been testing di-methyl ether (DME) as liquid propellant for a coaxial type pulsed plasma thruster (PPT). Liquid-PPTs can set arbitrary amount of ejected propellant at arbitrary timing, avoiding “particulate-emission” and “late-time ablation,” whereas solid propellant PPTs cannot. Additional feeding pressurant would be removed with the help of DME's expedient vapor pressure. We conducted a study to investigate the effect of capacitance on the thrust performance with our prototyped DME-PPT. The result showed that impulse bits increased monotonically with capacitor-stored energy E. In higher range of E, impulse bits Ibit in the case with capacitance C=3 μF was lower than in the case with the other Cs. The maximum Ibit and specific impulse Isp were 102 μNs and 575 s, respectively, at E=13 J (C=6 μF) and a mass shot Δm=18.4 μg in the range conducted in this study C=3–9 μF.

Optical measurements of unsteady phenomena on coaxial pulsed plasma thrusters

At Gifu University, to understand unsteady operation phenomena including plasma generation and exhaust processes on a coaxial pulsed plasma thruster (PPT), the luminescence behavior has been studied. In this study, spatial velocity distributions were compared with the results of electrostatic measurement and high-speed photography, and the applicability of a photosensor system to the plume measurements in the region slightly downstream of the nozzle was confirmed. Using the results of the measurements using the photosensor system, the dependence of the impulse bit of the PPT on the cavity geometry was investigated. In addition, direct observations of the ablation process in the cavity of a propellant with slits were carried out. The duration of late-time ablation measured for the propellant with slits is longer than that for the propellant without slits. The results of the observations reveal that the direct observation method is effective for investigating late-time ablation phenomena.

Characterization of Coaxial Pulsed Plasma Thruster by High-speed Photography

Pulsed plasma thrusters (PPTs) are electrothermal and electromagnetic thrusters that produce thrust in a discharge and have great potential as space engines owing to their simple structures. A coaxial PPT with solid propellant of cavity diameter of 3 mm and capacitor energy of 8 J has been designed and operated at Gifu University. To investigate unsteady phenomena of the PPT, optical measurements have been performed with photography using a high-speed camera. In past works, the luminescence from the PPT has been observed. The results show that the luminescence from the propellant continues after discharge. In the present study, to investigate the luminescence in detail, in addition to the photography of the luminescence without a filter, photographs were taken with a band-pass filter to obtain the luminescence from only the ionized gas, C+ for two different cavity lengths, 25 and 15 mm. By comparing the results between the unsteady behaviors of the luminescence with and without the filter, the luminescence phenomena were investigated. The luminescence from the propellant part without the filter in the last 34 and 15 μs for the cavity lengths of 25 and 15 mm was free of luminescence from C+. The time when the luminescence from C+ ends was close to the time of the end of discharge for each cavity length. These results suggests that the measurement of the luminescence from C+ is effective for understanding phenomena in the cavity during discharge.

Research and Development on Coaxial Pulsed Plasma Thruster with Propellant Feed

Research and Development on Coaxial Pulsed Plasma Thruster with Propellant Feed

Total impulse improvement of coaxial pulsed plasma thruster for small satellite

Research and development on pulsed plasma thruster (PPT) for small satellites has been conducted in Tokyo Metropolitan University (TMU). A PPT with coaxial electrode and PTFE cavity has been considered to gain larger impulse bit than that of parallel plate PPT. In this paper, the objectives are development of the PPTs with large impulse bit and large total impulse, to apply to a propulsion system for a small satellite. The electrodes and the propellant configuration were considered. As a result, divergent nozzle and larger cathode diameter improved thrust performances and larger impulse bit from 300 to 800 μNs was achieved. And the current reversal was appeared remarkable as its operation accumulated. Then, propellant-feeding mechanism was considered to improve its total impulse. With exchange of five propellants in each 20,000 shots, total impulse of 40 Ns was obtained. These results indicate that the PPT developed in TMU would provide various missions for a small satellite.