Air-breathing Electric Propulsion Research Articles

Aerodynamic Simulation Study of a Space Vehicle with Atmosphere-Breathing Electric Propulsion in a Free Molecular Gas Flow

Aerodynamic Simulation Study of a Space Vehicle with Atmosphere-Breathing Electric Propulsion in a Free Molecular Gas Flow

Direct simulation Monte Carlo-driven optimization of vacuum intakes for air-breathing electric thrusters in very low earth orbits

The performance of an atmosphere-breathing electric propulsion (ABEP) intake has been investigated with a focus on the direct simulation Monte Carlo (DSMC) method. A numerical dataset was derived from extensive DSMC analysis of rarefied flow across various intake configurations. The intake geometry, based on a concept from the literature, comprises a cylindrical body with four annular coaxial channels and a conical convergent diffuser. By maintaining the aspect ratio of the coaxial channels, the DSMC simulations were performed by changing three key parameters: inlet area, convergent diffuser angle, and operating discharge voltage, at altitudes ranging from 140 to 200 km. The analysis of the ABEP system revealed that altitude has the most significant influence on the discharge power, while the effects of the diffuser angle and inlet area are comparatively smaller. Analysis at fixed altitudes reveals a strong influence of altitude on discharge power, while the diffuser angle and the inlet area play a minor role. The results also show that the sensitivity of the discharge power to the diffuser angle increases as the altitude approaches the highest level of 200 km. Furthermore, an evolutionary-based optimization methodology was applied, taking into account the requirements of a drag-to-thrust ratio of less than 1 and a discharge power of less than 12 kW. Optimization analysis in the full altitude range revealed that the optimal diffuser angle falls within the narrow range of 15°–20°, corresponding to an optimal operating altitude range of 170–178 km.

Cathode-less RF plasma thruster design and optimisation for an atmosphere-breathing electric propulsion (ABEP) system

Atmosphere-breathing electric propulsion (ABEP) is a concept that ingests residual atmospheric gases as a source of propellant for an electric thruster, removing the need for onboard propellant storage. This would enable continuous low-thrust drag compensation, extending the lifetime of spacecraft in Very-Low Earth Orbit (VLEO); <250 km. VLEO is an appealing region for spacecraft operations, enabling new remote sensing missions with improved radiometric performance and spatial resolution, whilst reducing size, mass and power requirements, as well as mission cost. A preliminary design review and optimisation is therefore conducted for an ABEP system that uses the cathode-less radio frequency (RF) plasma thruster from Technology for Innovation & Propulsion (T4i) S.p.A. This removes the issue of thruster erosion by means of magnetic confinement and offers reduced susceptibility to varying atmospheric composition. A semi-empirical oxygen-nitrogen global source model (GSM) has been developed which considers the volume-averaged flux, momentum, and energy balance of the RF discharge. This includes a detailed chemistry model for the complex electron-molecular reactions and energy-loss channels of air plasma in the ionisation chamber. The GSM is coupled to an analytical model of flux balance for an air intake, verified by Direct Simulation Monte-Carlo (DSMC) simulation, to consider its design for maximum collection efficiency. This is then utilised in a robust multi-objective optimisation of the ABEP system, accounting also for spacecraft aerodynamics and power requirements.

Numerical simulations of a low-pressure electrodeless ion source intended for air-breathing electric propulsion

Abstract Air breathing electric propulsion (ABEP) systems offer a promising solution to extend the lifetime of very low earth orbit (VLEO) missions by using residual atmospheric particles as propellants. Such systems would operate in very low-pressure environments where plasma ignition and confinement prove challenging. In this contribution, we present results of a global plasma model (GPM) of a plasma ignited in a very low-pressure air mixture. The results are validated against experimental measurements acquired using a laboratory electrodeless ion source utilizing a resonator for plasma ignition. The device is specifically designed to operate within low-pressure environments as it holds potential applications in ABEP systems for VLEO missions. Parametric studies are carried out via GPM to investigate the resonant behavior and its implications. The potential of the model serving as a predictive tool is assessed through experimental validation against measured data, mainly investigating the extracted ion current dependency on operational pressure and external magnetic field strength. The verified model is further utilized to extrapolate additional information about the resonant plasma such as ion composition or a degree of ionization.

Design and optimization of a high thrust density air-breathing pulsed plasma thruster array

For satellites in Very Low Earth Orbit (VLEO) or the upper atmosphere, Air-Breathing Electric Propulsion (ABEP) can be used to counteract the drag effects of the atmosphere or provide attitude control without the need for onboard propellant storage. The Dense Plasma Focus (DPF) inspired the design of an ABEP Pulsed Plasma Thruster (TAMU PPT) which takes advantage of the simplicity, scalability, and high energy density of the DPF. This work details the design and testing of a PPT for use on satellites in the upper atmosphere. Testing was conducted at a variety of pressures and pulsing energies to determine which conditions provide maximum thrust and thrust to power ratios (TPR). It was found through single pulse experiments that a pressure of 6.3 hPa (corresponding to an altitude of about 35 km) and an energy per pulse of 9 J yield a TPR of 21 mN/kW. High frequency operation at 20 Hz in similar conditions showed a TPR of up to 28 mN/kW with a corresponding thrust of 4.8 mN. High speed imaging of the plasma plume found plasma velocities in excess of 2 km/s and identified metal ejecta with velocities up to 350 m/s. The pinching phenomenon associated with a DPF was also observed and a plasma jet velocity up to 50 km/s was recorded. The TPR of the TAMU PPT demonstrated in this work is comparable to that of other ABEP thrusters, however the TAMU PPT has an advantage when comparing thrust density. The TAMU PPT has a thrust density between 4.8 and 19 mN/cm2 while other ABEP thrusters have thrust densities between 0.019 and 0.36 mN/cm2.

On the critical parameters for feasibility and advantage of air-breathing electric propulsion systems

In recent years, air-breathing electric propulsion emerged as a potential enabling technology for long-duration space missions in Very Low Earth Orbit (VLEO). In this work, we show how the complex relation between mission environment, spacecraft configuration, and propulsive performance can be associated to a single requirement for full air-breathing drag compensation, which becomes less stringent for higher VLEO orbits. The impact of the spacecraft shape and size on the performed analysis is then evaluated and the results compared with conventional electric propulsion solutions with stored propellant, showing how the adoption of air-breathing propulsion becomes more advantageous, from a volume and mass fraction perspective, when the spacecraft scale is reduced.

Power transfer efficiency in an air-breathing radio frequency ion thruster

Abstract Due to a series of challenges such as low-orbit maintenance of satellites, the air-breathing electric propulsion has got widespread attention. Commonly, the radio frequency ion thruster is favored by low-orbit missions due to its high specific impulse and efficiency. In this paper, the power transfer efficiency of the radio frequency ion thruster with different gas composition is studied experimentally, which is obtained by measuring the radio frequency power and current of the antenna coil with and without discharge operation. The results show that increasing the turns of antenna coils can effectively improve the radio frequency power transfer efficiency, which is due to the improvement of Q factor. In pure N2 discharge, with the increase of radio frequency power, the radio frequency power transfer efficiency first rises rapidly and then exhibits a less steep increasing trend. The radio frequency power transfer efficiency increases with the increase of gas pressure at relatively high power, while declines rapidly at relatively low power. In N2/O2 discharge, increasing the N2 content at high power can improve the radio frequency power transfer efficiency, but the opposite was observed at low power. In order to give a better understanding of these trends, an analytic solution in limit cases is utilized, and a Langmuir probe was employed to measure the electron density. It is found that the evolution of radio frequency power transfer efficiency can be well explained by the variation of plasma resistance, which is related to the electron density and the effective electron collision frequency.

Monte Carlo Simulation of Inlet Flows in Atmosphere-Breathing Electric Propulsion

A comprehensive numerical study is performed to investigate gas flows inside the inlet of the atmosphere-breathing electric propulsion system using the direct simulation Monte Carlo method. The effects of inlet geometries and gas–surface interaction (GSI) accommodation coefficients on gas pressure and flow rate are analyzed in depth, and a comprehensive assessment of the compression and collection performances of inlets is conducted. Inlet geometries have noticeable effects on gas pressures inside inlets, and the gas pressures of the two concave inlets are the highest. Among the four inlets considered here, the rounded concave inlet performs the best in terms of compression and collection performances. Relative to the simple slope inlet, the rounded concave inlet achieves a rate of increase of over 120% in gas pressure and more than 110% in gas flow rate. GSI accommodation coefficients play a crucial role in both the pressure and mass flux inside the inlet, and the drop of the GSI accommodation coefficient from 1 to 0.2 brings about a considerable increase in the pressure and mass flux, achieving an increase of a factor of 8 and 5, respectively. Therefore, smoothing the inlet surface is an effective means of improving the compression and collection performances of inlets.

Deflagration thruster for air-breathing electric propulsion in very low Earth orbit

This study establishes the performance requirements for an air-breathing electric propulsion device to operate in very low Earth orbit (<450 km). We demonstrate that existing electric propulsion architectures, including electrostatic thrusters, fail to generate enough thrust while operating on flow rates of air that can be harvested in orbit. Instead, we develop and characterize a type of electromagnetic pulsed thruster that operates in a magneto-deflagration mode (i.e., forms an expansion wave during the acceleration process) that becomes more efficient as less propellant is used. We show this mode of operation can generate specific impulses up to 104 s, thrust per power >2 mN/kW, thrust densities >100 mN/m2, and thrust efficiencies up to 10% while consuming <100μg of air per discharge. We map these performance metrics on a candidate spacecraft to show the deflagration thruster can enable fully air-breathing drag compensation at altitudes ranging from ∼200 km to 350 km depending on its geometry and is extendable to other altitudes if stored propellant is utilized.

System design study of a VLEO satellite platform using the IRS RF helicon-based plasma thruster

To achieve a feasible lifetime of several years, most satellites are deployed in orbits higher than 400 km. Drag of residual atmosphere causes a slow orbit decay, resulting in the deorbit of the spacecraft. However, e.g. optical instruments or communication devices would significantly benefit from lower altitudes in the range of 150–250 km. A solution to achieve this could be the application of atmosphere-breathing electric propulsion (ABEP), where the residual atmosphere is used to generate continuous thrust that compensates the drag.Within the EU-funded DISCOVERER project, the Institute of Space Systems (IRS) developed an electrode-less RF Helicon-based Plasma Thruster (IPT) suitable for such applications. Ignition and preliminary discharge characterizations of the IPT have been carried out at IRS facilities, using argon, nitrogen and oxygen. To further characterize the plasma plume, a torsional pendulum has been designed to determine the (local) momentum flux in the plasma jet, as well as a three-axis magnetic B-dot probe to carry out time-varying magnetic field measurements. Various intake designs were investigated, opening the possibility to conduct studies on potential satellite platforms within the frame of the ESA-funded project RAM-CLEP.A design study for an Earth Observation and Telecommunication satellite operating at 150–250 km with an extended mission lifetime is currently being carried out. The first system assessment focused on the comparison of different spacecraft configurations (“slender body” and “flat body”) and intake designs (specular or diffuse) with regard to overall drag and ABEP performance requirements.In this contribution, the design approaches for the current thruster and the diagnostic methods are depicted. Moreover, the current status of the system assessment is presented. Upcoming experimental studies of the ABEP system e.g. within the ESA-project RAM-CLEP and additional activities planned on system assessment are outlined.

Plasma plume simulation of an atomic oxygen-fed ion thruster in very-low-earth-orbit

The plasma plume flow of an atomic oxygen-fed (AO-fed) ion thruster is numerically investigated as a simplification of the atmosphere-breathing electric propulsion (ABEP). A predictive analysis is conducted focusing on the ion backflow phenomenon and plume-background interaction in very-low-earth-orbit (VLEO). The computational framework employs two sequentially integrated numerical methods: a zero-dimensional (0-D) analytical model for the radio-frequency ion thruster and a hybrid method of the particle-in-cell (PIC) and direct simulation Monte Carlo (DSMC) techniques. The 0-D analytic model is employed for the prediction of exhaust conditions, while the hybrid PIC-DSMC method adopts these predictions to conduct the plasma plume simulations. A generalized collision cross-section model is introduced to enable consistent kinetic simulations for both AO and xenon propellants in VLEO atmosphere. The plasma plume simulations are conducted in an axisymmetric domain, including a cylindrical satellite body to consider wake flow. The exhaust ions exhibit diffusive transport transverse to the ion beam direction, implying the ion backflow. The backflowing ion current density can be increased in AO-fed thrusters, which require a high propellant flow rate to achieve a practical thrust. The AO-fed ion thruster shows a more active interaction between its plasma plume and the VLEO atmosphere compared to conventional xenon-based thrusters. The intensified plume-background interaction modifies the backflowing ion current density and the kinetic energy of individual ions, factors related to the spacecraft’s surface contamination.

Development and standalone testing of the Air-breathing Microwave Plasma CAThode (AMPCAT)

Air-breathing electric propulsion (ABEP) refers to a spacecraft in very-low Earth orbit (VLEO) harnessing upper atmospheric air as propellant for an electric thruster. This allows the orbital altitude to be maintained via drag-compensation, removing the need for on-board propellant storage and allowing a mission lifetime which is not limited by propellant capacity. A cathode (or neutraliser) is required for the high-specific impulse electrostatic thruster designs proposed for an ABEP application. One such study is the AETHER EU H2020 project, which aims to design an ABEP system that can be tested on-ground in a VLEO-representative environment. There is therefore a need to develop a cathode for ABEP as conventional thermionic hollow cathodes are susceptible to oxygen poisoning. The Air-breathing Microwave Plasma CAThode (AMPCAT) presented here is based on a plasma electron source, using a 2.45 GHz microwave antenna directly-inserted into the plasma volume to ionise neutral air particles. This study details the cathode design and the results of iterative standalone testing, with a particular focus on: (a) the identification of a dual-mode current emission, with transition from lower- to higher-current mode with air at bias values around 70 V between the extracting anode and internal cathode surfaces, (b) a comparison of performance relative to xenon, for which the peak extracted current is 30–40% higher than air at equivalent inputs, and (c) the effect of antenna electrical isolation, using alumina shielding thicknesses in the 0.1–0.7 mm range. Standalone cathode tests demonstrate 0.8 A of stable extracted current with 0.1 mg/s mass flow rate of a 0.48O2 + 0.52N2 mixture, relative bias of 80 V and input microwave power of 70 W. To the authors’ knowledge, the demonstration of an extracted current in the 1 A order using air, without visible material degradation after several hours of operation, is a novel development in the cathode literature.

Performance and plasma diagnostics of the Air-breathing Microwave Plasma CAThode (AMPCAT) coupled to a cylindrical Hall thruster

The Air-breathing Microwave Plasma CAThode (AMPCAT) has been developed for air-breathing electric propulsion in very-low Earth orbit. In this study, the standalone AMPCAT plasma characteristics are analyzed by means of several diagnostic tools and operation on xenon is compared to a conventional hollow cathode. A transition of AMPCAT extracted current from a lower (&amp;lt;0.1 A) to higher-current (&amp;gt;0.5 A) mode, triggered by increasing the negative cathode bias voltage, is accompanied by a significant rise in internal electron density and external electron temperature. The AMPCAT is coupled with a cylindrical Hall thruster in the 100–300 W power-level running on 0.5–0.7 mg/s of xenon, and the thrust is directly measured for cathode operation with both xenon and air. Stable thruster operation is demonstrated for the AMPCAT running on both propellants. For xenon, the performance is compared to a hollow cathode, which reveals matching discharge current profiles but a significantly higher thrust for the AMPCAT at low discharge voltages, approximately two times higher at 200 V. Langmuir probe measurements highlight a 30–40 V lower plasma potential in the cathode vicinity for the AMPCAT with xenon compared to both the hollow cathode and AMPCAT with air. This indicates a significantly improved coupling of cathode electrons to the thruster discharge, yielding an increased degree of ionization. Faraday probe and Wien filter results show that a larger current utilization efficiency drives the observed performance difference at low discharge voltages, rather than a significant change in ion acceleration or plume divergence.

Elliptical Orbit Design Based on Air-Breathing Electric Propulsion Technology in Very-Low Earth Orbit Space

Very-low Earth orbit (VLEO) space below 200 km is essential for high-quality communications and near-Earth space environment detection. Due to the significant atmospheric drag, orbital maintenance is required for spacecraft staying here. Based on air-breathing electric propulsion (ABEP) technology, this paper analyzed the orbital boundary conditions of the spacecraft under the constraints of parameters including slenderness ratio, thrust-to-power ratio, drag coefficient, and effective specific impulse. The energy balance is the key constraint for low VLEO orbits, which is determined by the drag coefficient, slenderness ratio, and thrust-to-power ratio. Under the existing technical conditions, the lowest circular orbit (along the terminator) is about 170 km. An elliptical orbital flight scheme is also analyzed to reach a 150 km perigee. A half-period control method was proposed based on the on–off control method for the elliptical orbit, which could enable the spacecraft to maintain a stable 150–250 km elliptical orbit.

Maintenance Strategy for Elliptical Orbit Satellite With Air-Breathing Electric Propulsion

Owing to the strong perturbation caused by atmospheric drag, it is inevitable to consume a significant amount of fuel for the orbital maintenance of ultralow Earth orbit (ULEO) satellites. Air-breathing electric propulsion (ABEP) has emerged at the right time, which makes use of the atmosphere to provide propulsion. Aiming to adopt an elliptical orbit, an ABEP-based satellite can collect the atmosphere near its perigee, which can be applied for propulsion to maintain the orbit shape. However, to guarantee propulsive efficiency, the thrusters should not work during the intake process, that is, near the perigee. This causes difficulties in orbital control methods because propulsion near the perigee is the most effective for simultaneous increasing the semi-major axis and eccentricity. To overcome this problem, this study proposes the “in-orbit balance” strategy, wherein the orbit maintenance can be achieved in closed orbits. Some of these orbits are used for air intake, while others are applied for propulsion near perigee. Considering the requirements regarding working gas, electric power, maneuverability and temperature, the constraints are analyzed for systematic parameter design. Using equivalent treatments, feasible satellite layout schemes can be searched for, and a constrained optimization problem is constructed to determine the solutions that satisfy the specific engineering requirements. Furthermore, optimal control for a continuous system is applied to allocate the thrust. The simulations verified the efficiency of the parameter design and confirmed that the ULEO can be controlled using the proposed strategy without resulting in extra fuel consumption.

Numerical analysis of inlet flows at different altitudes in the upper atmosphere

A comprehensive numerical study is performed to investigate gas flows inside the inlet of an atmosphere-breathing electric propulsion (ABEP) system operating in the upper atmosphere ranging from 120 to 300 km using the direct simulation Monte Carlo method. Gas pressure, mass flux, and aerodynamic drag are analyzed in depth in order to gain a deep understanding of the effects of operation altitude and the assumption of free molecular flow (FMF) on gas flows within the inlet. Computational results show that both the gas pressure and mass flux in the compression and ionization sections decrease with increasing altitude, indicating weaker compression and collection performances at higher altitudes. Therefore, careful attention should be paid to compression and collection performances of the inlet when it operates at higher altitudes. At altitudes smaller than 180 km, gas flows within the inlet are fully or partly characterized by transitional flows, so the FMF assumption tends to overestimate the gas pressure and underestimate the mass flux within the inlet resulting from the neglect of the collisions between the oncoming and reflected molecules. However, FMFs predominate within the inlet and even fill the entire inlet at altitudes larger than 180 km, so it is fairly reasonable to assume an FMF in the aerodynamic design of the inlet of an ABEP system.

Influence of applied magnetic field in an air-breathing microwave plasma cathode

The air-breathing electric propulsion concept refers to a spacecraft in very-low Earth orbit (VLEO) ingesting upper atmospheric air as propellant for an electric thruster. This compensates atmospheric drag and allows the spacecraft to maintain its orbital altitude, removing the need for on-board propellant storage and allowing an extended mission duration which is not limited by propellant exhaustion. There is a need for development of a robust, high current density and long life cathode (or neutralizer) for air-breathing electrostatic thrusters as conventional thermionic hollow cathodes are susceptible to oxygen poisoning. An Air-breathing Microwave Plasma CAThode is proposed to overcome this issue through the use of a microwave plasma discharge, producing an extracted current in the order of 1 A with 0.1 mg s−1 of air. In this paper, the effect of varying magnetic-field strength and topology is investigated by using an electromagnet coil, which reveals a significantly different behaviour for air compared to xenon. The extracted current with xenon increases by 3.9 times from the zero-field value up to a peak around 150 mT magnetic-field strength at the antenna, whereas an applied field does not increase the extracted current with air at nominal conditions. A non-zero magnetic-field with air is however beneficial for current extraction at reduced neutral densities. A distinct increase in extracted current is identified at low bias voltages with air for a field strength of around 50 mT at the internal microwave antenna, consistent across varying field topologies. The effect of a lowered magnetic-field strength in the orifice region is investigated through the use of a secondary coil, resulting in an extracted current increase of 25% for a relaxation from 6 mT to 1 mT, and demonstrating the beneficial impact of a locally reduced field strength on electron extraction.

Parametric study on the flight envelope of a radio-frequency ion thruster based atmosphere-breathing electric propulsion system

The atmosphere-breathing electric propulsion (ABEP) system utilizes atmospheric species as propellants to generate thrust for drag compensation of a satellite in very-low-Earth-orbit (VLEO). A parametric approach is used to assess the impact of different parameters on the flight envelope of a sample VLEO satellite with an ABEP system based on a radio-frequency ion thruster (RIT). The considered parameters include capture efficiency, maximum input power, solar activity, and atomic oxygen recombination factor. The NRLMSIS 2.0 atmosphere model is used to determine the flow conditions at VLEO, at a target altitude of 150–300 km with high, moderate, and low solar activity levels. DSMC method is employed to calculate the drag of the sample satellite with the ABEP system. The 0-D model of the RIT discharge chamber is used to predict the thrust of the RIT-based ABEP system. The flight envelope of a sample RIT-based ABEP system shows that drag can be compensated at altitudes between 196 km and 248 km. Increasing capture efficiency and maximum input power expands the feasible range for drag compensation to higher and lower altitudes, respectively. Also, the flight envelope shifts to higher altitudes with increasing solar activity levels. However, the atomic oxygen recombination factor of the intake device has minimal effects on the flight envelope.

Mechanism of capture section affecting an intake for atmosphere-breathing electric propulsion

Atmosphere-Breathing Electric Propulsion (ABEP) can compensate for lost momentum of spacecraft operating in Very Low Earth Orbit (VLEO) which has been widely concerned due to its excellent commercial potential. It is a key technology to improve the capture efficiency of intakes, which collect and compress the atmosphere for ABEP. In this paper, the mechanism of the capture section affecting capture efficiency is investigated by Test Particle Monte Carlo (TPMC) simulations with 3D intake models. The inner surface smoothness and average collision number are determined to be key factors affecting capture efficiency, and a negative effect growth model is accordingly established. When the inner surface smoothness is less than 0.2, the highest capture efficiency and its corresponding average collision number interval are independent of the capture section’s geometry and its mesh size. When the inner surface smoothness is higher than 0.2, the capture efficiency will decrease by installing any capture section. Based on the present results, the manufacturing process and material selection are suggested to be prioritized during the intake geometry design in engineering projects. Then, the highest capture efficiency can be achieved by adjusting the length and mesh size of the capture section.

Computational and experimental research on the performance of ECRIT ion source with nitrogen propellant

Electron cyclotron resonance ion thruster (ECRIT) with a diameter of 2 cm has the characteristics of no hot cathode and high specific impulse, which is suitable for the air-breathing electric propulsion system. In order to adapt to the atmospheric composition characteristics of nitrogen and oxygen in low orbit, the computational and experimental research on the performance of the ECRIT ion sourse with nitrogen propellant is an important basis for analyzing the feasibility of applying ECRIT to the air-breathing electric propulsion system. In this paper, the global model of the nitrogen ECRIT ion source with a diameter of 2 cm is established to calculate its performance. Then, the computational results are compared with the experimental results to analyze the difference. The research results show that when the input power of the ion source is 8 W and the gas flow rate is 2 ml/min, the computational and experimental results of the extracted ion beam current and thrust reach the maximum with the extracted beam current of 16.2 and 12.5 mA and the thrust of 476.6 and 368 μN, respectively. When the input power is 8 W and the gas flow rate is 0.6 ml/min, the computational and experimental results of the specific impulse are 2 095.8 and 1 855.6 s, both reaching the maximum value. The relative errors between the computational and experimental results of the extracted ion beam current, thrust and specific impulse all range from 2% to 32%. When the input power and gas flow rate used are 8 W and 1 ml/min in calculation, and 8 W and 0.8 ml/min in experiment, the ion source is on the optimal operating state. At this situation, the computational and experimental propellant utilization efficiencies with 17.8% and 16.2% respectively are high, and the ion energy loss with 443.9 and 596.2 W/A respectively is low.