Ablative Pulsed Plasma Thruster Research Articles

Side and front fast imaging of solid and liquid fed ablative pulsed plasma thruster’s discharge

Ultra-fast camera imaging is used to capture temporal evolution of the plasma discharge in an ablative Pulsed Plasma Thruster (PPT) fueled with solid polytetrafluoroethylene (PTFE) and liquid perfluoropolyether (PFPE), with the goal of comparing the mechanism of the expelled mass acceleration and getting insight into the physics of the processes involved in the discharge maintenance and impulse bit production. The fast photos are taken from two viewpoints (non-simultaneously), directing camera parallel (side view) and perpendicular (front view) to the ablative surface of the propellants, and are correlated with the relevant phases of the discharge current. Side view images reveal the presence of two distinct exhaust fractions, velocities of which differ by an order of magnitude. The speed of the faster fraction is comparable with the median of the ionized particles velocity expelled by the thruster. The vividness of the slow propagating component of the ablated mass, which mostly corresponds to neutral particles (based on velocity estimation) and lower mass bit of PFPE shots hint towards better propellant utilization by a liquid-fed PPT. Front view images confirm stable localization of PFPE discharge, forced by design of the propellant supplying plenum, whereas PTFE discharge is characterized by irregular path that covers an area of the propellant surface in random. The last half-period of the discharge is accompanied by vortices near the cathode spots. They are supposedly formed by evaporated electrode’s material and their trajectories imply further recombination in collisions with the electrodes, making this part of evaporated copper undetectable for charged particles diagnostics.

Pulsed plasma thruster exhaust reconstruction

The transient exhaust of an ablative pulsed plasma thruster is characterized experimentally for three capacitances and three discharge voltages, for discharge energies below 11 J. A novel analysis technique is introduced, which reconstructs the exhaust as a superposition of different Maxwellian-like ion groups. Each ion group is characterized by its own mean velocity, thermal spread, initial density, generation time and divergence rate. The time series of three probes working in the ion saturation regime are used to determine the model parameters by least-squares fitting. This approach allows a higher level of accuracy and insight than time-of-flight analysis based on direct feature correlation alone. A good fit of the main part of the discharge time series is achieved with just three ion groups, with mean velocities ranging in 50–70, 30–45 and 10–25 km s−1 respectively. Each ion group differs in the lateral divergence rate and axial thermal spread, and potentially corresponds with a different charge/mass ratio and/or creation time. Some trends with bank capacitance and discharge voltage are identified and discussed.

The Problem of Optimal Discharge Energy in an Ablative Pulsed Plasma Thruster

The Problem of Optimal Discharge Energy in an Ablative Pulsed Plasma Thruster

The Problem of Optimal Discharge Energy in an Ablative Pulsed Plasma Thruster

One promising trend in small-spacecraft development is the development of small electric propulsion systems (EPSs) based on ablative pulsed plasma thrusters (APPTs). The problem of optimal discharge energy in APPT that provides a minimal total mass of the EPSs has been considered. It has been shown that, at a given total impulse, discharge energy of an APPT has an optimal value that depends on the specific energy intensity of power capacitors, the specific thrust impulse of the engine, mass of electronic units, and other elements of the propulsion system circuit. It has been concluded that the calculation of the optimal discharge energy allows reducing the total mass of the propulsion system during the design of an APPT-based EPSs.

Discharge Characteristics and System Performance of the Ablative Pulsed Plasma Thruster with Different Structural Parameters

Under the given initial discharge energy level, altering the electrode structural parameters of the Ablative Pulse Plasma Thruster (APPT) is an effective way to improve the performance of the thruster. The purpose of this study is to reveal the underlying mechanism of the effect of changing the electrode structure parameters on the performance of the APPT system and to offer targeted support for researchers to optimize the design of APPT structure. With rectangular and tongue-shaped electrode configurations at various electrode flare angles, electrode lengths, and electrode spacings, the discharge characteristics, propellant ablation characteristics, and thruster performance of the APPT are systematically investigated. The underlying mechanism of how changing the electrode’s configuration parameter affects the performance of the thruster is identified by fitting and predicting the parameters of the APPT discharge circuit and system performance under various operating conditions. The results show that using tongue-shaped electrodes is more effective than using rectangular electrodes in terms of enhancing the inductive gradient of the electrodes, transferring more energy to the discharge channel, and increasing the squared integral value of the discharge current. As a result, the tongue-shaped electrode APPT performs better than the APPT with rectangular electrodes, as a consequence. The thruster’s performance can be enhanced for the same electrode configuration by increasing the electrode flare angle within a certain angle range; however, the improvement is extremely limited. Additionally, in the case of small electrode spacing, increasing the electrode flare angle can enhance the thruster’s performance more effectively.

Influence of Different Energy Supply Methods on Performance of Ablative Pulsed Plasma Thrusters

For a certain configuration of ablative pulsed plasma thruster, changing the initial energy is an effective way to optimize the performance of the thruster. Different energy supply methods will not only affect the macro discharge characteristics and system performance of the thruster, but also affect the micro characteristics of the plasma in the discharge channel, the equivalent parameters of the discharge circuit, energy conversion efficiency and so on. In this study, the discharge characteristics, ablation characteristics and plasma motion characteristics of the thruster under different energy supply methods are analyzed experimentally and theoretically. The results show that using the method which includes increasing capacitance to heighten the initial energy of system under the same voltage, using a low-voltage, large-capacitor power supply method under the same initial energy, can effectively increase the impulse generated by Lorentz force and its proportion in the impulse bit. Moreover, the proportion of the ablative propellant effectively ionized and accelerated by Lorentz force increases. Therefore, the thruster has higher specific impulse and efficiency.

Development and testing of low-power EP based on the ablative pulsed plasma thruster for an ERS SSC

The paper presents the appearance of a low-power electric propulsion system for a small remote sensing satellites based on an ablative pulsed plasma thruster. Preliminary analysis for the purpose and fields of application of ablative pulsed plasma thrusters (APPT) was carried out. The design appearance of a thruster for small remote sensing satellites was identified, and the main directions of APPT further development were defined. Methods for calculating the APPT propulsion characteristics were developed. The breadboard model of APPT for small remote sensing satellites was developed and made (APPT-350). Research tests of the breadboard model of APPT for Earth remote sensing small spacecrafts were carried out.

Development of power processing unit for a low-power ablative pulsed plasma thruster

Research Institute of Applied Mechanics and Electrodynamics of the Moscow Aviation Institute (RIAME MAI) develops ablative pulsed plasma thrusters with low power consumption for small spacecraft. The paper considers the development of power processing unit (PPU) for such thrusters. PPU under discussion should meet the following requirements: design simplicity, small overall dimensions, ensuring the required power and frequency of thruster operation and high energy efficiency.

A reliable data-driven model for Ablative Pulsed Plasma Thruster

Ablative Pulsed Plasma Thrusters (APPTs) are high specific impulse electric space propulsion system, but a reliable model equivalent of the experimental model is still unavailable. In this paper, a reliable model is developed based on APPT experimental data by using Machine Learning (ML) ecosystem. The goals of this study are to justify the accuracy and reliability of the newly built APPT model with the existing experimental and simulation model. For four sets of operating conditions, 600 experimental and simulation test operations are done. The experimental voltages and currents are measured with a high-voltage probe and a Rogowski coil, respectively. The simulation voltages and currents are gathered by running the respective simulation program. Comparison results show that the newly built APPT model has better accuracy and reliability than the simulated APPT model as compared to real APPT used in the experiment. This data-driven approach provides a novel way of designing a reliable alternative model of physical APPTs.

Impulse and Performance Measurements of Electric Solid Propellant in a Laboratory Electrothermal Ablation-Fed Pulsed Plasma Thruster

Electric solid propellants are advanced solid chemical rocket propellants that can be controlled (ignited, throttled and extinguished) through the application and removal of an electric current. This behavior may enable the propellant to be used in multimode propulsion systems utilizing the ablative pulsed plasma thruster. The performance of an electric solid propellant operating in an electrothermal ablation-fed pulsed plasma thruster was investigated using an inverted pendulum micro-newton thrust stand. The impulse bit and specific impulse of the device using the electric solid propellant were measured for short-duration test runs of 100 pulses and longer-duration runs to end-of-life, at energy levels of 5, 10, 15 and 20 J. Also, the device was operated using the current state-of-the-art ablation-fed pulsed plasma thruster propellant, polytetrafluoroethylene (PTFE). Impulse bit measurements for PTFE indicate 100 ± 20 µN-s at an initial energy level of 5 J, which increases linearly with energy by approximately 30 µN-s/J. Within the error of the experiment, measurements of the impulse bit for the electric solid propellant are identical to PTFE. Specific impulse when operating on PTFE is calculated to be about 450 s. It is demonstrated that a surface layer in the hygroscopic electric solid propellant is rapidly ablated over the first few discharges of the device, which decreases the average specific impulse relative to the traditional polytetrafluoroethylene propellant. Correcting these data by subtracting the early discharge ablation mass loss measurements yields a corrected electric solid propellant specific impulse of approximately 300 s.

Experimental investigation on the plasma morphology of ablative pulsed plasma thruster with tongue-shaped and flared electrodes

Ablative pulsed plasma thrusters (APPTs) are considered as an attractive propulsion option for station-keeping and drag makeup purposes for mass- and power-limited satellites. In order to understand the physical mechanism of APPTs, high-speed camera and optical emission spectroscopy are utilized to investigate the plasma characteristics including the spatial distribution and composition between the electrodes. The plume images and spectra at different times and positions are experimentally recorded, and the spatial distribution, composition, and trajectory of plasmas can be concluded through analyzing them. With the increase of the distance from the ablation surface, two clusters of plasmas near the anode and cathode meet downstream, and the species and density of plasmas tend to be uniform.

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.

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.

Thermodynamic Properties of Hydroxylammonium Nitrate-Based Electric Solid Propellant Plasma

Electric solid propellants are advanced solid chemical rocket propellants controlled by electric current. An electric solid propellant may also be used in an electric propulsion system: specifically, an ablative pulsed plasma thruster. Previous experiments with the electric solid propellant known as high-performance electric propellant, or HIPEP, suggest its ablation processes are similar to the traditional propellant polytetrafluoroethylene (). Better understanding of the ablation and the resulting propulsion performance of HIPEP requires a model of its vapor composition, bulk plasma quantities, and thermodynamic properties. Such a model was developed and favorably compared with previous analogous model predictions found in the literature for . The electric solid propellant vapor composition was predicted in the temperature range of 500–40,000 K at 1 bar pressure. Low temperatures (less than 2000 K) were dominated by water (), carbon dioxide (), and molecular nitrogen (), with results at 700 K matching to within 10% of the previous combustion model predictions. At high temperatures (greater than 25,000 K), the vapor was strongly ionized and dominated by carbon (), oxygen (), nitrogen (), and hydrogen () ions. The calculated enthalpy and specific heat were higher for high performance electric propellant (HIPEP) vapor than , suggesting increased thermal losses could be inherent to ablation-controlled HIPEP discharges.

An Ablative Pulsed Plasma Thruster Based on Multianode Electrode Geometry

This article presents a special surface discharge achieved by a multi-anode electrode geometry (MAEG). This surface discharge can be applied to the ablative pulsed plasma thruster (APPT) and solve the problem of beam divergence. The MAEG changes the discharge characteristics of the surface discharge and extends the $Z$ -pinch existing within the thruster to the plasma plume outside of the nozzle compared with the conventional electrode geometry (CEG). The enhanced-MAEG further promotes the confinement effect of plasma plume by the $\theta $ -pinch. The discharge phenomenon shows that the visible plasma plume lengths of the MAEG and the enhanced-MAEG have been prolonged effectively compared with that of the CEG. The electron densities of the MAEG and the enhanced-MAEG in axial direction have also been measured to reach 3.5 and 4 times of that of the CEG. Through microthrust measurement, the MAEG APPT and the enhanced-MAEG APPT have been proven to have superior performance in terms of the impulse bit and the thrust-to-power ratio than that of the CEG APPT. The MAEG and the enhanced-MAEG could achieve improved propulsion performance with its system advantages by special electrode geometries. It has been found that the MAEG and the enhanced-MAEG have the potential to be used in the field of microsatellites propulsion.

Mechanisms for the Formation of Parasitic Propellant Consumption in an Ablative Pulsed Plasma Thruster

Results of spectroscopic studies of electric discharge stages in the channel of an ablative pulsed plasma thruster (APPT) are presented. The initial stage has been shown to affect further discharge development. Shock surface ionization of the propellant has been found to be the dominant mechanism for “parasitic” propellant consumption of an APPT. The results of experimental studies of the dependence of propulsion efficiency and consumption on the distance between operating surfaces of APPT propellant bars are given.

Electric Solid Propellant Ablation in an Arc Discharge

Electric solid propellants are advanced solid chemical rocket propellants that can be controlled (ignited, throttled, and extinguished) through the application and removal of an electric current. Electric solid propellants are also being considered for pulsed arc ablation electric thrusters, such as the pulsed plasma thruster. The focus of this work is the electrical and ablation characteristics of electric solid propellant within an arc discharge. Arc discharges of 5–20 J per pulse were created within a cylindrical cavity, and results for the electric solid propellant are compared with polytetrafluoroethylene (PTFE), which is a traditional propellant in ablative pulsed plasma thrusters. The data indicate that the electric solid propellant has higher specific ablation per pulse () relative to PTFE (), which quantitatively agrees with an ablation energy balance model. For both propellants, the equivalent circuit resistance and inductance of the plasma arc are 50 mΩ and 125 nH, respectively. Analyses are presented indicating that the physics of propellant ablation is similar for both propellants with the differences in the observed specific ablation owing to differences in the thermodynamic properties between propellants.

A review of the characterization and optimization of ablative pulsed plasma thrusters

The increasing number of small satellites has resulted in greater consideration of electric propulsion (EP) as potential thruster systems. Of these, ablative pulsed plasma thrusters (APPTs) are a type of EP system that is suitable for microsatellites. They have recently attracted attention for application in small satellites due to their simple and compact system. However, from a scientific perspective, we still have limited understanding regarding APPT mechanisms such as the acceleration mechanism, plasma propagation, and associated physical phenomena. From an engineering perspective, widespread adoption can be aided by the optimization of APPTs, including improvements in the thrust efficiency, energy bank miniaturization, and a better understanding of the plasma plume characteristics and possible interactions with spacecraft. These are some issues that still remain as challenges in the research and development of APPTs. This review discusses our current understanding of the fundamental mechanisms behind APPT operation as well as some new insights and recent progress associated with APPTs. Typical characterization methods for APPTs are presented along with discussion of the operation process and relevant physical phenomena. To support an engineering perspective, we also discuss optimization approaches for APPTs such as changes in the thruster geometry, alternative propellants, and electrical circuit optimization. New insights and novel approaches resulting from the use of modern technology in APPT research are also presented, along with current understanding that attempt to tie scientific perspectives with engineering observations, such as the physical reasons behind certain performance enhancements from optimization efforts. Finally, some remaining challenging issues are also discussed.

Impact of a High-Power Pulsed Plasma Flow with a Surface of High-Temperature Materials

The objective of this work is to study the behavior of surface layers of high-temperature metals in their interaction with a powerful pulsed plasma flow produced by a high-power ablative pulsed plasma thruster. This plasma generator produces plasma flows with a directed velocity of (7–9) × 106 cm/s, an initial diameter of 1.5–2 cm, and a maximum number density of about 1018 cm–3, as well as a maximum power of 5 GW. The main measured values are the residual temperature of the tungsten specimens and the evaporated mass. Also, metallographic analysis of the specimens was performed. The basis of the research method is to analyze the experimental data with the help of a numerical model describing the heating and evaporation of the material upon absorption of pulsed energy fluxes taking into account the evaporation kinetics based on the Hertz–Knudsen expression. Based on the developed numerical model and the obtained experimental data, the kinetics of evaporation of tungsten at high power fluxes to the surface (up to 1 GW/cm2) is investigated.

A new ablation model for ablative pulsed plasma thrusters

Pulsed plasma thrusters have been widely used all over the world in the past several decades. A new and effective model to calculate the ablated mass bit (the mass ablated from the propellant in each discharge) of ablative pulsed plasma thrusters is presented in this paper. The model is based on the random degradation of molecular chains in polytetrafluoroethylene (PTFE) and statistical methods. New expressions are obtained for the ablated PTFE mass bit as a function of the energy used in PTFE ablation, the ratio of the energy used to successfully break the carbon-carbon bonds to the energy used in PTFE ablation and the ratio of the mass that breaks away from the ablation surface of the PTFE to the total ablation mass. The introduction of statistical methods allows us to define a simplified model that is useful for the prediction of the ablated mass bit and for preliminary design guidance. The obtained theoretical results for the mass bit show very good agreement with the results from experimental studies. The range of the average relative error is approximately −15.7% to 12.4%. Finally, this new model can be used to evaluate the ablated mass bit, in addition to predicting the lifetime and guiding the design of new ablative pulsed plasma thrusters.