Electrothermal Thruster Research Articles

Computational study of a helium-propellant microwave electrothermal thruster

The microwave electrothermal thruster (MET) utilizes wave-exited microdischarges to heat gas flows, enhancing the specific impulse of the thruster. Our computational study investigates a 17.5 GHz helium-propellant MET, employing a two-dimensional, axisymmetric fluid model of plasma coupled with electromagnetic wave and gas flows. The discharges operate in the glow regime, remaining weakly ionized, and in thermal non-equilibrium. The plasma densities reach approximately 1020m−3, and the gas temperature is around 2000 K. Even a slight off-resonant frequency operation results in a significantly lower plasma density and gas temperature. Gas heating, primarily driven by electromagnetic Joule heating, plays a critical role in influencing the overall discharge behavior. The measured peak thrust and specific impulse are 8.24 mN and 292 s, respectively, at a mass flow rate of 3.2 mg/s with 30 W of power. Compared to a cold gas thruster, there is a significant increase in the specific impulse by a factor of approximately 1.7. The enhanced performance trades off with propulsive efficiency, which decreases by a factor of 1.5 from the peak 65% at 10 W. This is due to higher energy losses to cavity walls from heat conduction with increased power. These findings underscore the critical balance between the input power and mass flow rate to improve the MET performance, highlighting the importance of power management to maximize thrust and efficiency. Furthermore, the predicted thrust and specific impulse agree well with experimental values for nominally similar MET thruster studies in the literature.

Microwave electrothermal thruster with surface wave plasma

AbstractA new microwave electrothermal thruster (MET) with small dimensions (length of 25 mm and diameter of 6 mm), utilizing surface wave (SW) plasma at low average power consumption Pav ˜ 2 W, is made. It uses microwaves at a frequency of 2.45 GHz for plasma generation within a dielectric chamber enclosed in a metal shield. The SW is excited at the plasma‐dielectric interface by a coaxial launcher and a standing wave is created due to the wave reflection from the shield. The high‐density plasma produced in this resonant chamber heats up the argon gas, which expands through the nozzle in the axial direction. The temperature of the gas decreases from 1900 K at atmospheric pressure to 1200 K at 10 Torr ambient pressure. This small‐sized MET shows stable low‐pressure operation with a thrust of 4 mN and an efficiency of 8% and is suitable for nanosatellite orbit control.

Differential Analysis of GPX Traces on Historical and Recent Satellite Imagery

High Power rocketry had observed great achievement in past few decades. It has a large tailing of propulsion technologies which are not only effective but also sustainable. Among all the propulsion types, like the most widely used in earlier era, chemical propulsion, Electric propulsion had gained there roots in the field due to its advantages in performance and durability in space missions. In this paper we will be dealing with the types of electric propulsion arrangement and network with its potentiality applications. The paper starts from giving an insight of traditional propulsion ways, like chemical propulsion, which were commonly used in rockets. With which the key features of electric propulsion including electrostatic, electrothermal, electromagnetic and electrodynamic thrusters have been discussed. With witch the main working principles, features and applications are discussed. Also the limitations of earlier propulsion system like efficiency of fuel and thrust have been discussed. In addition to which the the challenges faced during the era of the change in electric propulsion have been discussed which will help us to get an intuition of the phase change. In this paper we can come across at the result which demonstrates the advantage of electric propulsion over the old methods of propulsion. The advantages of electric Propulsion like extended mission duration, which increase the efficiency of mission, and the expanded payload capacity which gives a wide range to mission portrait. To sumup with which we will find the advantages of electric propulsion over other types of propulsion, with the advantage of electric propulsion and its wide features which gives a lead to our missions. Keywords: propulsion systems, electric propulsion, high-powered rockets, Hall Effect thrusters, ion thrusters, Specific Impulse, hybrid propulsion, experimental validation, cost-benefit analysis, space applications.

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.

Bidirectional vortex stabilization of a supersonic inductively coupled plasma torch

Radio-frequency (RF) inductively coupled plasma (ICP) torches using a supersonic nozzle have many industrial materials processing applications and have also been proposed as novel electrothermal plasma thrusters for space propulsion. The gas injection method in plasma torches plays an important role in both gas heating dynamics and overall discharge stabilization. Here, we investigate reverse vortex gas injection into a supersonic ICP torch for RF powers up to 1 kW, argon mass flow rates between 15 and 180 mg s−1, and plasma torch pressures from ∼270 Pa to ∼50 kPa. In this configuration, gas is injected tangentially just upstream of the nozzle inlet. This produces a bidirectional vortex flow field where gas first spirals upwards along the outer edge of the plasma torch walls, before then reversing direction at the torch end and spiralling back down through the central plasma region towards the nozzle exit. Results are compared to a more conventional forward vortex configuration where gas is instead injected tangentially from the upstream end of the torch, and which forms a unidirectional vortex that spirals towards the downstream nozzle. While performance is similar for gas flows below 80 mg s−1, we show that at higher mass flow rates both the effective torch stagnation temperature and thermal efficiency can be increased by almost 50% with reverse vortex injection. Considering that the measured RF antenna-plasma power transfer efficiency is similar for both configurations, this enhancement occurs because of the unique bidirectional vortex flow field which leads to reduced gas-wall heat losses and consequently an increased enthalpy flow leaving the torch.

Design and Construction of a 2.45 GHz Microwave Electrothermal Thruster

A microwave electrothermal thruster (MET thruster) has been constructed, consisting of a microwave plasma chamber 12 cm long and 8 cm in inner diameter and a micronozzle 10 mm long with 1 mm in diameter. The microwave plasma is produced by 2.45 GHz microwave frequency at a power of 1 kW, and the feed gas is Ar at a pressure of 10^-3mTorr. Microwave energy is transmitted into the cavity and electrons sre connected to the wave's electric field. Thus, the electrons are accelerated by microwave electric fields. Microwave plasma discharge is formed based on the interaction of electrons with neutral gas particles. Then, the plasma acts as a resistive load and absorption of microwave energy, raises the temperature of the gas or plasma. Gas heating increases the gas pressure and is released through the nozzle. The plasma density and electron temperature are 2.35×〖10〗^17 m^(-3) and 1.2 eV, respectively. The thrust and specific impulse are 10 mN and 100 s

Experimental investigation of electric propulsion systems using C12A7 electride hollow cathodes

The development and experimental investigation of two low-power electric propulsion concepts using compact heaterless C12A7 electride (C12A7:e-) hollow cathodes is presented. The first concept represents an electrothermal thruster, in which a cathode discharge is used to heat a gas that is subsequently accelerated in a nozzle-shaped anode. The second propulsion system is an attempt to develop a sub-500 W magnetoplasmadynamic thruster (MPDT) that uses a rectangular discharge channel that allows to increase the applied magnetic field and thus lower the necessary discharge current. Extensive parameter studies with both concepts were conducted, and the thrust and discharge properties of different geometric and operational configurations were determined. This work is a follow-up publication of a previous paper (Gondol and Tajmar in CEAS Space J 14:65–77, 2021).

Исследования проектного облика малого космического аппарата с электротермической двигательной установкой

The paper presents the results of solving the multivariate problem of forming the design of a small spacecraft with various propulsion systems based on freons with electrothermal thrusters. As initial data, the required reserves of characteristic velocities for maintaining the active lifetime for circular orbits of 350-390 km for small spacecraft with different midsection areas are determined. Specific and dimensional-mass characteristics of propulsion systems of small spacecraft are determined using a set of basic design parameters of a small spacecraft obtained by random search. Ranked rows of alternative variants of small spacecraft have been formed as the complex total criterion decreases by the multicriteria method of prioritization for various principles of designing a small spacecraft with a propulsion system.

Publisher’s Note: “Additive-manufactured single-piece thin multi-layer tungsten heater for an electrothermal thruster” [Rev. Sci. Instrum. 92, 114501 (2021)

Publisher’s Note: “Additive-manufactured single-piece thin multi-layer tungsten heater for an electrothermal thruster” [Rev. Sci. Instrum. 92, 114501 (2021)

Additive-manufactured single-piece thin multi-layer tungsten heater for an electrothermal thruster.

In this study, a novel single-piece thin multi-layer tungsten resistive heater was successfully fabricated using additive manufacturing and tested as an electrothermal thruster. The heater has 12 resistive layers, with each layer having a thickness and height of 0.15 and 81mm, respectively, and can provide high heating efficiency. A single-piece or monolithic heater was manufactured via additive manufacturing technique, which drastically improved its reliability and decreased its manufacturing cost. In the heating and thrust measurement tests that used nitrogen gas as a propellant, the heater reached a gas temperature of ∼2000K at a 140-A heater current without experiencing any failure. The tungsten-heater resistance linearly increased with an increase in temperature due to the temperature dependence of tungsten's resistivity. The specific impulse and thrust increased with the heater temperature in accordance with the theoretical prediction. Even including a voltage drop due to a contact resistance, the achieved heater efficiency reached 63% at a 100-A heater current even without a thermal insulation around the thruster. The heater efficiency decreased with an increase in the heater temperature due to heat loss to the surroundings. The heat-loss analysis indicated that both thermal conduction and radiation heat losses were crucial for improving the heater performance at a high-temperature operation of over 2000K.

Characterization of a radio-frequency inductively coupled electrothermal plasma thruster

A radio-frequency (RF) inductively coupled electrothermal plasma thruster operating with argon is experimentally characterized for different mass flow rates, RF powers, and propellant injection configurations. Depending on the propellant mass flow rate, significant neutral gas heating is observed with effective stagnation temperatures around 2000 K (giving a maximum estimated thrust and specific impulse of about 100 mN and 125 s, respectively) for absorbed powers between 300 and 500 W. A self-consistent theoretical discharge model is developed and used to study the basic physics and operation of RF electrothermal thrusters, and predictions of the gas temperature are in good agreement with experimental measurements. The model identifies primary power inefficiencies as electron-neutral excitation losses and neutral gas heat losses to the thruster walls. Both experimental and theoretical results indicate that a relatively high stagnation pressure (of the order of 100 Torr or higher) is critical for high performance. For pressures significantly below this the electron-neutral collisional power transfer is too low to effectively heat the neutral gas.

Design of a hollow cathode thruster: concepts, parameter study and initial test results

Two electric propulsion concepts have been developed at Technische Universität Dresden as spin-off devices of regular hollow cathodes and initial testing has been conducted. Both devices represent millinewton thrusters that take advantage of thermionic electron emission using the low work function materials C12A7, LaB6, and thoriated tungsten in different design configurations. The first concept represents an electrothermal thruster which generates thrust by expanding and accelerating a heated propellant in a nozzle. Initial thrust measurement tests were carried out which showed thrust levels well above cold gas thrust, but low thrust efficiencies. The influence of different geometric parameters on the discharge properties and the performance is investigated and presented. The second thruster concept is a novel electromagnetic device in which charge carriers in a plasma discharge are accelerated by an applied magnetic field that is orthogonally oriented to the discharge current. Initial tests with C12A7 were not successful, but the functionality of the concept was shown by thrust measurements using a thoriated tungsten wire as an electron emitter.

Numerical investigation of the plasma processes for propellant heating in electrothermal plasma thruster for nanosatellites

AbstractNumerical investigation of the plasma processes in a cylindrical chamber with small dimensions of a novel microwave electrothermal plasma thruster for nanosatellites has been conducted. The absorbed microwave power from the electrons in the plasma column of the surface wave discharge is included in the computational model as a heat source with Gaussian distribution. The computational model takes into account the elastic and inelastic collisions of the electrons with the atoms in the ground state and two excited states (−s, −p) and the processes of recombination and deactivation of the plasma species in the volume and on the walls of the chamber. The computational model includes the flow of neutral gas and the processes in the plasma for effective heating of neutral particles by collisions not only with electrons but also with ions. Selected combinations of input power and propellant mass flow rates are used as initial parameters for the numerical investigation. The results show that at higher mass flow rates the heating of the neutral gas is more effective and at power levels of 4 W and propellant mass flow rate of 3 mg/s the electrothermal plasma thruster demonstrates effective performance and thrust levels in the order of 1 mN.

Sulphur-fuelled Surface Arc Thruster for Propelling Nanosatellites

Pulsed electrothermal thrusters use kiloampere discharge currents for polytetrafluoroethylene (PTFE) ablation and sublimation. However, higher arc currents cause higher electromagnetic interference (EMI). A 10 J surface arc thruster (SAT), which adopts current regulating diodes (CRD), was developed that enables significant reduction in EMI. A CRD limits the discharge currents to 5 A in spite of the applied voltage. A low-melting-point sulphur propellant has been used that enables low-discharge currents to efficiently ablate it and accelerate it electrothermally. In this paper, the near-term potential for elemental sulphur propellant in SAT is investigated. The advantages of sulphur with respect to PTFE are presented. First, we measured the pulse width of the main discharge. Then sulphur propellant proved superior to PTFE propellant in discharge duration. The mean pulse widths of PTFE and sulphur are 3.38 and 22.1 ms, respectively. Second, we measured the pressure rise in the vacuum chamber after each discharge. The mean pressure rises of PTFE, sulphur powder and sulphur solid are 0.43, 0.94 and 1.9 mPa, respectively. Sulphur powder experienced the least intensity of discoloration in comparison with other propellants. The discharge initiator misfired frequently during PTFE ablation, but it ignited successively during sulphur ablation. Experimental results indicate sulphur is a suitable propellant for surface discharge propulsion, and for low arc currents is superior to PTFE.

Performance Impacts of Metal Additive Manufacturing of Very Small Nozzles

AbstractFor small satellite electrothermal or chemical thrusters, nozzle throat diameters may be less than 1 mm. At these sizes, the effects of flow surface roughness on nozzle performance must be ...

Numerical study of plasma-fluid characteristics and thrust performance of a low-power argon inductively coupled plasma electrothermal thruster

Low-power (from the sub-kilowatt range up to a few kilowatts) inductively coupled plasma (ICP) electrothermal thrusters for space propulsion are potential alternatives to low-power arcjet thrusters, which are often implemented on geostationary satellites for north–south station keeping. We develop an axisymmetric two-dimensional magnetohydrodynamic (MHD) numerical simulation technique using a two-temperature plasma model as a computer-aided engineering tool for low-power argon ICP electrothermal thrusters. Furthermore, to the best of our knowledge, this is the first study to provide a basic understanding of the plasma-fluid characteristics of low-power ICP electrothermal thrusters. To this end, we perform the MHD numerical simulation for a low-power argon ICP electrothermal thruster model that was developed for thrust measurement experiments to validate the proposed numerical simulation technique. The numerical results indicate that the plasma flow produced in the low-power argon ICP electrothermal thruster model is basically in a strongly thermal and ionizing nonequililbrium state. In addition, the numerical results show that the experimentally measured thrust forces can be reproduced with an underestimation of 10% or less over the RF input power range of 0–0.6 kW considered in the experiment. Moreover, the numerical results suggest that the main reason why the performance of the low-power argon ICP electrothermal thruster model is significantly lower than the target performance is the considerable heat loss of more than 90% of the input power through the flow-channel wall.

Inductively coupled plasmas could be a low-power alternative for satellite propulsion

Pending improvements in thrust performance, simulations suggest inductively coupled plasma electrothermal thrusters may be suitable for atmospheric electric propulsion systems.

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.

Feedback control of a spacecraft electro-dynamic thruster

In this paper, the possibility of feedback servo control of a spacecraft thruster's specific impulse by a boundary feedback system is theoretically considered. The motivation to introduce feedback control is two-fold. The first is to stabilize any inherent plasma instabilities and the second is regulate the output specific impulse of the thruster. Two cases are considered: Electro-thermal thrusters and Electro-dynamic thrusters. The chamber inlet temperature in the case of the electro-thermal thruster or the boundary electric field potential responsible for generation of the velocity of the plasma ions in the case of an electrodynamic thruster, are controlled by feedback so as to regulate the thruster's specific impulse. By introducing typical disturbances in the plasma ionization voltage, it is shown using a two-dimensional fluid model and a suitable boundary feedback law, where the chamber inlet temperature or applied boundary potential is proportional to the error in the specific impulse and the desired specific impulse, that the specific impulse of the thruster may be regulated and held constant. The robustness of the control system is numerically tested, by a two dimensional simulation model using McCormacks's method. The Navier-Stokes equations, including the magneto-hydrodynamic variables were discretized and simulated, using the explicit MacCormack method for a typical nozzle domain. The numerical results for the open and closed-loop velocity fields were obtained and the specific impulse was computed from these fields. It was thus shown that not only the stability of the plasma is realized but also that the specific impulse is regulated as desired.

Methods and means of the experimental research of the electrothermal arcjet thrusters of spacecrafts

The subject of the research in this article is the measurement of the main characteristics of the electrothermal arcjet thrusters in the process of the working fluid acceleration during the bench tests. The purpose is to form a bench research facility to study the characteristics of the electric heating propulsion system for use in a space tugboat, taking into account the launch dynamics, operating conditions and shutdown dynamics. Tasks: measuring the thrust of the electric propulsion system as the main characteristic in the established and transient mode; determining the flow rate of the working fluid correlated with the thrust; measuring the parameters of electrical consumption ‒ currents, voltage, power in static and dynamic modes; measuring the external conditions of the simulating space environment, namely the vacuum, in the test chamber. The methods used are: tensometry, differential pressure gauge, four-probe measurement of power supplied to engine heaters, steam generator and tank of the electric heating system, ionization-thermocouple vacuum measurement. The following results have been obtained: the composition of the information and control system consisting of the tensometric sensor used with the electric heating propulsion system, flowmeter of the working body of the electric heating propulsion system, power meter of the heater, steam generator and the tank of the electric heating propulsion system, ionization-thermocouple sensors for vacuum determination. The software product for graduation of the tensometric sensor has been developed. The scientific innovation of the obtained results is as follows: the method of measuring the thrust of the electric heating propulsion system by means of tensometric sensor has been further developed, the method of measuring the flow of the working body of the electric heating propulsion system has been further developed, the method of measuring the power of the engine heater, steam generator, tank of the electric heating propulsion system has been further developed. For the first time an algorithm for graduation of the tensometric sensor for the electric heating propulsion system was developed and implemented in the software product.