Pulsed Plasma Thruster Performance Research Articles

Performance of pulsed plasma thruster at low discharge energy

As the size of satellites scales down, low-power and compact propulsion systems such as the pulsed plasma thruster (PPT) are needed for stabilizing these miniature satellites in orbit. Most PPT systems are operated at 2 J or more of discharge energy. In this work, the performance of a PPT with a side-fed, tongue-flared electrode configuration operated within a lower discharge energy range of 0.5‒2.5 J has been investigated. Ablation and charring of the polytetrafluoroethylene propellant surface were analyzed through field-effect scanning electron microscopy imaging and energy-dispersive X-ray spectroscopy. When the discharge energy fell below 2 J, inconsistencies occurred in the specific impulse and the thrust efficiency due to the measurement of the low mass bit. At energy 2 J, the performance parameters are compared with other PPT systems of similar configuration and discussed in depth.

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 the discharge circuit inductance on the ablative pulsed plasma thruster performance

The most promising option for motion control of small spacecrafts with weight less than 50 kg is a pulsed plasma thruster with an energy of less than 20 J (micro-PPT), which has such advantages as low weight, structure simplicity, precise pulse control, sufficiently long life time, constant operational readiness and low inertia. It was believed that PPT with low discharge energy (less than 20 J) needed reduction of the initial inductance, despite the fact that the low capacity of the capacitor bank would not allow to achieve the optimal C/Lo ratio. However, the results of calculations and experiments show that a slight increase in the initial inductance of the discharge circuit, contrary to the assumptions, has a positive effect on the micro-PPT characteristics, reducing the propellant consumption and mainly increasing the specific impulse without significant reduction of the thrust impulse (unit impulse). This feature of the discharge circuit of the PPT with a low discharge energy allows us to discover new solutions in micro-PPT designing and improve the thruster specific characteristics.

Numerical Modeling of Pulsed Plasma Thruster Performance with Teflon Ablation and Ionization

A numerical model employing the ablation rate and ionization of Teflon is developed to analyze pulsed plasma thruster (PPT) performance. A one-dimensional lumped circuit analysis model is implemented to calculate the discharge current. The Saha equation and Keidar Teflon ablation model are used to analyze the ionization of carbon and fluorine molecules produced by the Teflon gas. The numerical model also predicts the electron temperature distribution and simplified plasma shape during the discharge. Numerical simulation results with different voltage conditions compare reasonably well with experimental data. The Teflon gas flux reaches its maximum value at the maximum electron temperature and then gradually decays. The multivalent ions are investigated for PPT operation.

Physical mechanisms and factors influencing inductive pulsed plasma thruster performance: a numerical study using an extended magnetohydrodynamic model

The inductive pulsed plasma thruster (IPPT) is a sort of spacecraft propulsion device. To promote the development of IPPTs, the working process and performance influencing factors are investigated in this study, using a novel magnetohydrodynamic model, with an exterior circuit coupled and extended magnetic field involved in the computation.By exploring the spatial distributions of different physical fields and the global properties of the thruster, a comprehensive physical picture of the operation process is illustrated, which is comprised of the structure evolution of the plasma current sheet, the relationship between thrust force and circuit current, the coupling effect that derives from plasma back to the circuit, and the energy deposition and conversion process of the system. Besides, the positive contribution of the secondary current sheet to the ultimate thrust performance is also confirmed.Further investigations discuss the impacts of varied factors including the operating parameters, initial gas distributions, and metal structures nearby. The results indicate that by adjusting propellant mass to achieve identical specific energy, the specific impulse and efficiency of the thruster can remain almost constant under varied discharge energy. Both compressing the initial gas against the face of coil, and ameliorating its radial uniformity, can help enhance the coupling between the plasma and drive-coil. Conductive structures nearby, especially the grounded metal plate that mounts the capacitors, can impair the acceleration of the plasma, unless it is located out of the electromagnetic coupling distance of the drive-coil.This study may help to reveal the underlying physical mechanisms and improve the designs of IPPTs.

Optimization of the Energy Distribution in Ablative Pulsed Plasma Thrusters

Regarding space missions, electric propulsion efficiency is a hot topic due to the development of microsatellites. However, ablative pulsed plasma thrusters, which were the first electric propulsion devices, suffer from low efficiency (less than 10%). In this study, a double-discharge ablative pulsed plasma thruster that allocates the total energy to two capacitors and discharges at different locations was designed to improve the efficiency of these thrusters. The performance of the double-discharge ablative pulsed plasma thruster was calculated based on experimental results under four different working conditions. The results showed that the double-discharge ablative pulsed plasma thruster performance was significantly improved relative to that of a regular ablative pulsed plasma thruster, in which the total energy is stored in a single capacitor. When the total energy remains constant (2.25 J), as the energy allocated to the secondary capacitor increases, the thrust efficiency and specific impulse increased significantly, respectively. Finally, theoretical analysis combined with the experimental results revealed the successful separation of the ablation–ionization and acceleration processes for the double-discharge ablative pulsed plasma thruster, contributing to the improvement of its performance. This research provides a reference for the design of high-performance ablative pulsed plasma thrusters.

An Investigation of Discharge Characteristics of an Electrothermal Pulsed Plasma Thruster

The capillary-based pulsed plasma thruster (PPT) has high-energy transfer efficiency, which provides potential to improve traditional PPT performance. In this paper, the discharge characteristics of a capillary-based PPT have been investigated. First, the typical process of voltage and current waveform is recognized. It is featured with a fast voltage drop (~25 ns) and a consequent damped sine wave oscillation discharge. Then based on the circuit analysis, the equivalent plasma resistance has been calculated. The current waveform is well fit with a linearly increasing resistance model. The influence of charging voltage has also been studied. The deposited energy increases while the transfer efficiency decreases with the increasing applied voltage. The cavity dimension has effects on the equivalent circuit parameters. The results show that longer cavity and smaller diameter lead to more deposited energy and higher transfer efficiency. The influence of voltage and capillary dimension on the mass shot characteristic has been discussed. Finally, the estimated impulse bit has been calculated and the impulse bit characteristics have been discussed. These results can contribute to the further optimization of PPT.

Theoretical Modeling of Pulsed Plasma Thruster Performance with Teflon Ablation

A performance analysis for a pulsed plasma thruster(PPT) has been conducted to predict the thrust and current change. Two models were implemented - a one-dimensional lumped circuit analysis model and the Teflon ablation model provided by Michael Keidar. The circuit model incorporating resistance and inductance models was adapted to predict the magnitude of the discharge current. Numerical simulations like current discharge rates with different voltages were reasonably well compared with experimental data. The effects of Teflon ablation on thruster characteristics were investigated.

Single- and Repetitive-Pulse Conical Theta-Pinch Inductive Pulsed Plasma Thruster Performance

Impulse bits produced by conical theta-pinch inductive pulsed plasma thrusters possessing cone angles of 20°, 38°, and 60°, were quantified for 500-J/pulse operation by direct measurement using a hanging pendulum thrust stand. All three cone angles were tested in single-pulse mode, with the 38° model producing the highest impulse bits at roughly 1-mN-s operating on both argon and xenon propellants. A capacitor charging system, assembled to support repetitively pulsed thruster operation, permitted testing of the 38° thruster at a repetition rate of 5 Hz at power levels of 0.9, 1.6, and 2.5 kW. For similar conditions, the average thrust measured during repetitive-pulse operation exceeded the value obtained when the single-pulse impulse bit is multiplied by the repetition rate, suggesting that a greater impulse bit per pulse was produced when operating in the repetitive-pulse mode.

Pulsed plasma thruster performance for miniaturised electrode configurations and low energy operation

A one-dimensional electromechanical model based on accurate calculations of inductance and a detailed description of the inhomogeneous magnetic field distribution which accelerates the plasma discharge produced by a miniaturised pulsed plasma thruster ( μ PPT ) has been developed. The analytical model relates μ PPT performance to electrode geometry parameters and circuit parameters and shows good agreement with experimental results. The theoretical study has shown that flared electrode geometries enhance performance by optimising the inductance gradient and hence the energy available to accelerate the plasma sheet. Thruster performance has been experimentally characterised for a discharge energy range of 1–8 J for propellant surface areas of 0.15–1.5 cm 2. The performance of parallel electrode configurations has been shown to display a strong dependence on aspect ratio relative to that of flared geometries. Rectangular flared electrodes produce impulse bits considerably higher than parallel geometries under the same operating conditions. Results show that for configurations where aspect ratio may influence performance increases in electrode separation rather than decreases in electrode width produce enhanced performance for the same propellant surface area.

The effect of easily ionized elements Na and K on the performance of pulsed plasma thruster using water propellant

In view of the low thrust power ratio caused by the high resistance of pulsed plasma thruster using water propellant, the paper argues that the easily ionized elements Na and K with low ionic potentials are added in the water propellant to improve its performance. The measurement of the discharging current and plasma emission spectrographic analysis prove the improvement. The experiments show that the elements Na and K have certain effect on the improvement of the performance of pulsed plasma thruster: In comparison with water propellant, the NaCl and KCl water propellant has a lower total resistance and a higher ratio of thruster power and specific impulse, and the NaCl water propellant has a slightly stronger effect on pulsed plasma thruster than the KCl. The plasma emission spectrographic analysis is in consistent with the experiment of measuring the discharging current: The elements Na and K can intensify the plasma emission spectrographic signal.

Effect of ceramic nozzle on performance of pulsed plasma thruster

Pulsed plasma thruster (PPT) is one of the most promising electric propulsions system. Its main drawback is the low thrust efficiency. A theoretical analysis is given in this paper for exploring the reasons. The low electrothermal acceleration efficiency is concluded as an important factor. For improving the electrothermal acceleration, a new electrode structure associated with a nozzle was designed and adopted. The experimental result shows that the new electrode impaired the electromagnetic acceleration slightly. But the overall thrust was improved significantly. It means that the electrothermal acceleration was enhanced distinctly. The thrust efficiency was also improved remarkably. The simple electrode structure with a nozzle implicates an efficient way to improve the performance of PPT.

A study on the efficiency of pulsed plasma thruster (PPT)

PurposeThe pulsed plasma thruster (PPT) is one sort of promising electric thrusters, but the low efficiency is always a big problem for PPT. Many researches have been working on solving this problem. However, there is still no significant breakthrough. This paper aimed to discuss some methods for improving the PPT's efficiency.Design/methodology/approachThe causes for the low efficiency by analyzing the theoretical models of PPT are discussed, and the influences on PPT's performance investigated in terms of the structural or electrical parameters.FindingsThe change of structural or electrical parameters which influence the thrust efficiency significantly are taken into account. In the process of designing PPT, optimized structural and electrical parameters should be chosen to attain a better performance.Research limitations/implicationsThis analysis is mainly based on the parallel‐rail PPT.Practical implicationsThe result of analysis is adopted, the higher thrust efficiency of PPT is expectable.Originality/valueBy introducing theoretical models and analyzing results of some researcher's experiments, this paper offers a method of designing PPT and attempts to help designer to choose appropriate structural and electrical parameters.

Effect of Propellant Temperature on Efficiency in the Pulsed Plasma Thruster

Abstract : A pulsed plasma thruster (PPT) benefits from the inherent engineering simplicity-and reduced tankage fraction gained by storing the propellant as a solid. The solid is converted to the gaseous state and accelerated by an electric discharge across the propellant face. Previous research has concluded that as little as 10% of the consumed propellant is converted to plasma and efficiently accelerated. The remaining propellant is consumed in the form of late-time vaporization and particulate emission, creating minimal thrust. Critical to improving the PPT performance is improving the propellant utilization. The present work demonstrates one possible method of increasing the PPT propellant efficiency. By measuring the PPT thrust, propellant consumption, and propellant temperature while varying the power level, duration of the experimental run, and total propellant mass, a correlation is established between decreased propropellant temperature and increased propellant efficiency. The method is demonstrated by performance measurements at 60 W and S W, which show a 25% increase in thrust efficiency, while the propellant temperature decreases from 135 to 42 deg C. Larger increases in the efficiency may be realized on-orbit where operating temperatures are commonly subzero. The dependence of propellant consumption on temperature also creates systematic errors in laboratory measurements with short experimental runs, and orbit analyses where the PPT performance measured at one power level is linearly scaled to the power available on the spacecraft.

Life Extension Strategies for Shuttle-Deployed Small Satellites Using Pulsed Plasma Thrusters

Abstract : At typical Space Shuttle altitudes atmospheric drag is the dominant force limiting satellite on-orbit life (typically <100 days). The pulsed plasma thruster (PPT) is ideally suited to extend the life of small satellites deployed from the Shuttle due to its low system mass and volume, high specific impulse, and inert solid propellant (Teflon). The objective of this study was to identify and analyze life extension strategies for Space Shuttle-deployed small satellites using the pulsed plasma thruster. A generalized analysis is presented which is applicable to a broad range of satellite, PPT performance and life characteristics. Many PPT thrusting strategies were identified, enough to fit most spacecraft operational scenarios, for significantly extending small satellite on-orbit life. Within the limits of typical small satellite power to mass ratios, the most capable of these strategies, designated Lift & Coast, requires the least amount of propellant and is capable of extending life 1-2 years with state-of-the-art PPT technology. In addition, within the context of PPT operation, preferential launch windows and strategies for reducing satellite drag were also discussed.