Electrothermal Thruster Research Articles

Current-Free Electric Double Layer in a Small Collisional Plasma Thruster Nozzle Simulation

A computational fluid dynamics and plasma model of a collisional (∼ a few Torr) radiofrequency (at 13.56 MHz) argon plasma capacitively coupled in a converging-diverging nozzle (applied to the optimisa- tion of electrothermal plasma thrusters for space use) shows the formation of a strong stationary current-free double layer (CFDL) at the 1.5 mm diameter nozzle throat for a downstream pressure of ∼ 0.1 Torr. The cycle average magnitude of the double layer potential is ∆ΦDL = 77 V and the electron temperature at the high potential edge of the double layer is kBTe = 2.64eV, yielding a strength of ∆ΦDL/(kBTe) ∼ 30. The double layer is 1.2 mm wide which corresponds to ∼ 90 Debye lengths. The axial electric field of the double layer accelerates ions along the nozzle to a maximum drift velocity of 17 km s−1, about 3.3 times the ion sound speed, and their kinetic energy is transferred to neutrals by ion-neutral charge exchange col- lisions. The ion transit time τi through the potential structure spontaneously forming at the nozzle throat is about 5 times the radiofrequency excitation period τRF. These findings are discussed in the broader context of double layer physics and the dynamics of their formation as well as in the context of electrothermal thruster optimisation in which neutral propellant heating via ion-neutral charge exchange collisions is the main source of thrust.

Global model of microwave plasma assisted N2O dissociation for monopropellant propulsion

A global numerical model is developed to study the capability of a Microwave Electro-thermal thruster (MET) to operate with Nitrous Oxide gas (N2O), for spacecraft monopropellant thruster applications. N2O is a “green” propellant with interesting properties but highly inert when used as monopropellant; the non-equilibrium activation effect of the plasma is able to promote combustion of the gas. The model is self-consistent and features different temperatures for the vibrational modes of the mixture in order to account for possible vibrational activation effects of the plasma; a detailed kinetic scheme is presented, comprising rates for vibrational energy exchanges between the modes. The model shows that good theoretical efficiencies can be achieved by operating the MET with N2O, with thrust to power ratios up to 1 mN/W and specific impulse up to 200 s. The main channels of plasma assisted dissociation are examined, and the role of vibrational non-equilibrium is addressed.

Plume diagnostics of BUSTLab microwave electrothermal thruster using Langmuir and Faraday probes

This study presents the Langmuir and Faraday probe measurements conducted to determine the plume characteristics of the BUSTLab microwave electrothermal thruster (MET). The thruster, designed to operate at 2.45 GHz frequency, is run with helium, argon and nitrogen gases as the propellant. For the measurements, the propellant volume flow rate and the delivered microwave power levels are varied. Experiments with nitrogen gas revealed certain operation regimes where a very luminous plume is observed. With the use of in-house-built Langmuir probes and a Faraday probe with guard ring, thruster plume electron temperature, plasma density and ion current density values are measured, and the results are presented. The measurements show that MET thruster plume effects on spacecraft will likely be similar to those of the arcjet plume. It is observed that the measured plume ion flux levels are very low for the high volume flow rates used for the operation of this thruster.

About energy efficion of the electrothermal thruster with additive heat for the small spacecraft

About energy efficion of the electrothermal thruster with additive heat for the small spacecraft

Synthesis of Diamonds from the Microwave Plasma with the Use of Supersonic Gas Flows

A new method of deposition of diamond films with the use of supersonic gas flows activated by a microwave discharge is implemented for the first time. The operation principle of the proposed gas-discharge system is similar to that of a microwave electrothermal thruster. A mixture of hydrocarbons and hydrogen is used as a plasma-forming gas. It is demonstrated that the proposed method allows the plasma-forming gas to be used at pressures far above the upper limit of the pressure range of modern microwave plasma systems for chemical vapor deposition of diamond films (approximately 40 000 Pa).

Spatio-temporal plasma heating mechanisms in a radio frequency electrothermal microthruster

Low-power micro-propulsion sources are currently being developed for a variety of space missions. Electrothermal plasma thrusters are of specific interest since they enable spatial control of the power deposition to the propellant gas. Understanding the mechanisms whereby electrical power is coupled to the propellant will allow for optimization of the heating and fuel efficiencies of electrothermal sources. Previous studies of radio frequency (RF) plasmas have shown a dependence of the gas and electron heating mechanisms on the local collisionality. This is of particular importance to thrusters due to the large pressure gradients that exist between the inlet and outlet when expanding into vacuum. In this work, phase-resolved optical emission spectroscopy and numerical simulations were employed to study plasma heating in an asymmetric RF (13.56 MHz) electrothermal microthruster operating in argon between 186–226 Pa (1.4–1.7 Torr) plenum pressure, and between 130–450 V (0.2–5 W). Three distinct peaks in the phase-resolved Ar(2p1) electron impact excitation rate were observed, arising from sheath collapse heating, sheath expansion heating, and heating via secondary electron collisions. These experimental findings were corroborated with the results of two-dimensional fluid/Monte Carlo simulations performed using the Hybrid Plasma Equipment Model (HPEM). The influence of each mechanism with respect to the position within the plasma source during an α-γ mode transition, where plasma heating is driven via bulk and sheath heating, respectively, was investigated. Sheath dynamics were found to dictate the electron heating at the inlet and outlet, this is distinct from the center of the thruster where interactions of secondary electrons were found to be the dominant electron heating mechanism. Optimization of the heating mechanisms that contribute to the effective exhaust temperature will directly benefit electrothermal thrusters used on miniaturized satellite platforms.

Efficiency Estimation of Electrothermal Thrusters Use in the Control System of the Technological Spacecraft Motion

The influence of various control systems of the orbital motion of a technological spacecraft on the level of microacceleration of its internal environment is simulated. Conclusions are drawn about the effectiveness of control systems with different actuators for realization of certain gravitationally sensitive processes onboard a spacecraft.

Direct Thrust Measurements of an 8-GHz Microwave Electrothermal Thruster

Direct thrust measurements of an 8-GHz microwave electrothermal thruster were performed. Experimental work involved the redesign of the thruster cavity to minimize its weight and the construction of a vertical thrust stand. Thrust measurements were made using ammonia and simulated hydrazine (50% and 100% decomposed) as propellants. Testing was performed over a range of 100–250 W of microwave power input into the thruster. For all three propellants, the thrust increased with the increasing mass flow at higher microwave power levels able to sustain plasma at higher flow rates. Ammonia specific impulse increased with increasing cavity input power, reaching 264 s at 250 W. The 50% and 100% simulated decomposed hydrazine exhibited an initial increase in specific impulse with increasing power; however, further increases in power did not result in increased specific impulse. Higher levels of thrust and specific impulse at the same microwave input power were measured for the 100% decomposed simulated hydrazine compared to the 50% decomposed due to the absence of ammonia in the 100% decomposed hydrazine.

Research and Development of Electrothermal Pulsed Plasma Thruster Systems Onboard the Osaka Institute of Technology 2nd PROITERES Nano-Satellite

The Project of Osaka Institute of Technology Electric-Rocket-Engine onboard Small Space Ship (PROITERES) was started at Osaka Institute of Technology in 2007. After successfully launching the 1st PROITERES nano-satellite, the 2nd PROITERES nano-satellite with electrothermal Pulsed Plasma Thrusters (PPTs) has been developed since 2010. The main mission is powered flight with very long distance, i.e. changing more than 10km in altitude on near-earth orbits by high-power PPTs. In this research, a 30W-class PPT system was developed. The PPT achieved the maximum impulse bit of 2.47mNs, specific impulse of 342s and total impulse of 90.1Ns with repetitive 72,000 shots. Repetitive 100,000 shots could be also carried out using a newly developed Power Processing Unit (PPU). Furthermore, a long-time operation system with a Multi-Discharge-Room type PPT (MDR-PPT) head was developed. In the present paper, we explain development status of EM-PPT system including MDR-PPT, PPU, and capacitor.

Global Energy Transfer Model of a Microwave Electrothermal Thruster Operating With Helium Propellant at 2.45-GHz Frequency

Microwave electrothermal thruster (MET) is a type of space propulsion system that uses free floating plasma to heat the propellant gas. This paper presents a 0-D model that investigates the energy transfer from the electromagnetic wave to the propellant gas via atmospheric pressure plasma, and evaluates the steady-state plasma properties of an MET thruster. In the model, governing equations for electron particle balance, electron energy, balance and heavy particle energy balance equations are solved to obtain the plasma parameters. Electron temperatures of around 1 eV, electron number densities of $10^{19}~\#/\text {m}^{3}$ , and heavy particle temperatures of about 3000 K are evaluated for the free floating plasma inside the MET resonant cavity for 1200-W delivered power at 2.45-GHz frequency. Obtained solutions are in good agreement with the experimental results from the literature. Along with the plasma parameters, rocket performance parameters, and specific impulse and thrust are computed. The developed model is also used to predict the plasma and performance parameters of a prototype MET system developed at the Bogazici University Space Technologies Laboratory.

Gas dynamic model of electrothermal thrusters of small spacecraft and possibility of applying microwave heating of a working

This paper is devoted to development and approbation of the gas dynamic model of ammonia thruster with low power consumption and ultra small thrust for picosatellite weighing up to 5 kg and possibility of applying microwave heating of a working fluid. It is shown, that simplest electrothermal thruster consisting of propellant tank, solenoid valve, expension cavity and heating chamber can provide ultra small trust due to gas dynamic processes and small heat supply. The results of the study set tasks for further design of small spacecrafts microwave generators.

Numerical investigation of the electric field distribution and the power deposition in the resonant cavity of a microwave electrothermal thruster

Microwave electrothermal thruster (MET), an in-space propulsion concept, uses an electromagnetic resonant cavity as a heating chamber. In a MET system, electromagnetic energy is converted to thermal energy via a free floating plasma inside a resonant cavity. To optimize the power deposition inside the cavity, the factors that affect the electric field distribution and the resonance conditions must be accounted for. For MET thrusters, the length of the cavity, the dielectric plate that separates the plasma zone from the antenna, the antenna length and the formation of a free floating plasma have direct effects on the electromagnetic wave transmission and thus the power deposition. MET systems can be tuned by adjusting the lengths of the cavity or the antenna. This study presents the results of a 2-D axis symmetric model for the investigation of the effects of cavity length, antenna length, separation plate thickness, as well as the presence of free floating plasma on the power absorption. Specifically, electric field distribution inside the resonant cavity is calculated for a prototype MET system developed at the Bogazici University Space Technologies Laboratory. Simulations are conducted for a cavity fed with a constant power input of 1 kW at 2.45 GHz using COMSOL Multiphysics commercial software. Calculations are performed for maximum plasma electron densities ranging from 1019 to 1021 #/m3. It is determined that the optimum antenna length changes with changing plasma density. The calculations show that over 95% of the delivered power can be deposited to the plasma when the system is tuned by adjusting the cavity length.

Nonequilibrium modeling study on plasma flow features in a low-power nitrogen/hydrogen arcjet thruster

Gasdynamic flow features in an electrothermal arcjet thruster with a mixture of 1:2 nitrogen/hydrogen as the working gas have been studied by a two-temperature numerical simulation. Seven species and 17 kinetic processes are included in the chemical kinetic model used to represent dissociation, ionization, and the corresponding recombination reactions in this nitrogen/hydrogen mixture system. Based on the gas flow characteristics inside the arcjet nozzle, a new method is introduced to define the edge of the cold boundary layer, which is more convenient to analyze the evolution and development of plasma flow in an arcjet thruster. The results show that the arcjet thruster performance is determined largely by the exchange of energy and momentum between the low-density, high-temperature arc region and the high-density, cool-flow region near the nozzle wall. A significant thermal nonequilibrium is found in the cold boundary layer in the expansion portion of the nozzle. The important chemical kinetic processes determining the distribution of hydrogen and nitrogen species in different flow regions are presented. It has been shown that the reaction rate of hydrogen species ionization impacted by electrons is much higher than that of nitrogen species ionization in the center of the constrictor of the arcjet thruster. This indicates that hydrogen species is very important in the conversion of applied electric energy into thermal energy in the constrictor region of the arcjet thruster.

Investigation of Variation in the Performance of an Electro Thermal Thruster with Aerospike Nozzle

One of the most recently developed modes of propulsion is electric propulsion. The commonly used chemical propulsion systems have the advantage of a high Specific Impulse as compared to that of ion propulsion systems. However, owing to the efficacy of ion propulsion systems, it is considered the future of space exploration.Electro thermal thrusters produce thrust by using electrical fields to force hot plasma out of the nozzle with certain exit velocity. The plasma’s exit velocity and the system’s thrust capacity, as of now, are insufficient for space travel to be conducted within a reasonable time. I intend to study the possibility of improving the thruster’s performance by using an aerospike nozzle as an exit nozzle which meets the conditions required for the thruster to function appropriately. I shall be studying the plasma plume exit velocity variation with respect to the nozzles used. Also, a thermal analysis will be conducted in order to find the correct material for the nozzle.

A Short Review of Experimental and Computational Diagnostics for Radiofrequency Plasma Micro-thrusters

Experimental and computational diagnostics for radiofrequency plasma micro-thrusters are presented, based on the low power (10–100 W) electrothermal thruster prototype, Mini Pocket Rocket, developed for use on the Cubesat nanosatellite platform. Computer simulations include computer fluid dynamics simulations and particle in cell simulations while experimental results are obtained using a variety of electrostatic, optical and momentum probes. The output and limitations of each diagnostic are discussed within the context of device development for space use.

Nanosecond optical imaging spectroscopy of an electrothermal radiofrequency plasma thruster plume

Nanosecond optical imaging spectroscopy is employed to investigate the spatio-temporal dynamics of the plasma plume expanding from a 4.2 mm-diameter, 20 mm-long cylindrical capacitively coupled electrothermal radiofrequency (rf) driven thruster using 10 W of power at 12.50 MHz and an argon pressure of 1.5 Torr. On-axis, the plume exhibits four distinct peaks of optical emission intensity within the rf period. The plume has a spherical shape with a transient radial extension (during half of the rf cycle) at the thruster exit plane due to an rf current to ground when the grounded electrode acts as an anode.

Deposition of Fluorinated Diamond-Like-Carbon Films by Exposure of Electrothermal Pulsed Plasmas

Thin amorphous carbon films are deposited on silicon substrates by exposure to pulsed plasmas where the feed gas is mainly generated from the ablation of an insulator. An electrothermal pulsed plasma thruster with a discharge room in an insulator rod is used as the pulsed plasma for the ablation of the insulator, and the material of the insulator rod is poly(tetrafluoroethylene) (PTFE). The pulsed plasma, in which the estimated electron density is on the order of 1022–1023 m-3, is generated by the stored energy in the capacitor. The deposition rate, which depends on the stored energy, is lower than 1 nm per pulse in our experiment. The maximum hardness measured using a nanoindenter is about 7 GPa at a stored energy of about 2.7 J, beyond which the hardness of the films decreases with the increase in stored energy. Raman spectroscopy is also carried out to examine the formation of fluorinated diamond-like carbon films. In addition, the influence of dilution gas on the properties of the deposited films is also investigated.

S192023 大阪工業大学・超小型人工衛星プロイテレスのシステム設計と構造系開発([S19202]小型宇宙システム(2))

The 15kg-class nano-satellite "PROITERES" (The Project of Osaka Institute of Technology Electric-Rocket-Engine onboard Small Space Ship) with electrothermal pulsed plasma thrusters (PPTs) and high resolution camera is being developed in Osaka Institute of Technology. The mission of this satellite is powered flight of nano-satellite by PPTs in orbital and observation of Kansai district in Japan with a high-resolution camera. PROITERES will be launched from Satish Dhawan Space Centre (India) to the sun-synchronous orbit of 670 km in 2011.We developed Flight Model (FM) of the satellite, including electrothermal PPT system, high-resolution camera system, onboard computer system, communication system and grand station, electric power system, attitude control system etc, in 2010-2011. In this paper, we report all systems of PROITERES and its structure.

Research and Development of the Pulsed Plasma Rocket Engine System onboard Osaka Institute of Technology Micro Artificial Satellite

The Project of Osaka Institute of Technology Electric-Rocket-Engine onboard Small Space Ship (PROITERES) was started at Osaka Institute of Technology in 2007. In PROITERES, a micro satellite with electrothermal pulsed plasma thrusters (PPTs) will be launched in 2010. The main mission is the first powered flight of micro satellite by electric thruster all over the world. This study aims at improvement in performance by changing configuration of PPTs. The total impulse of about 5 Ns was achieved with a teflon cylindrical discharge chamber 9.0 mm in length and 1.0 mm in diameter in 53,000-shot operation with 2.43 J/shot. Finally, the engineering model of PPT system was developed, and it is under operation as final test.

S1901-1-6 大阪工業大学・電気推進ロケットエンジン搭載超小型人工衛星プロイテレスの開発(小型宇宙システム(1),社会変革を技術で廻す機械工学)

The Project of Osaka Institute of Technology Electric-Rocket-Engine onboard Small Space Ship (PROITERES) was started at Osaka Institute of Technology in 2007. In PROITERES, a nano-satellite with electrothermal pulsed plasma thrusters (PPTs) will be launched in 2010. The main mission is to achieve powered flight of nano-satellite by an electric thruster and to observe Kansai district in Japan with a high-resolution camera. The flight model of the satellite was successfully developed, and the satellite FM was under final test.