Thruster Operation Research Articles

Enhancing underwater thruster anomaly detection with support vector glow encoding description

As the driving system of autonomous underwater vehicles (AUVs), the healthy operation of underwater thrusters is crucial to ensure the performance of AUVs. Due to the challenge of obtaining faulty samples during normal operation, anomaly detection of underwater thrusters is often carried out using only normal samples. To address this, Support Vector Glow Encoding Description (SVGED) is proposed for anomaly detection of underwater thrusters trained with only normal samples. Specifically, the deep features of the operational vibration signals collected from the underwater thruster are extracted by a deep convolutional autoencoder based on a glow model. The glow-encoded data are then used to identify the boundary of normal operation for the underwater thruster. When a new sample enters the trained SVGED model, it can be effectively identified as representing a normal or anomalous condition of the underwater thruster. The developed SVGED was evaluated using AUV experiments. When using the proposed SVGED method, the accuracy reaches 99.81%, Results show that the proposed method can effectively real time detect anomalies in the underwater thruster compared to other methods. It lays the foundation for ensuring the healthy operation of AUVs.

A numerical model calibrated by in-situ measurement for predicting the underwater radiated noise of an azimuth thruster structure of a passenger ship

In a global context of limiting the impact of noise pollution on the behavior of marine fauna, the understanding of acoustic radiation phenomena from underwater sound sources of a passenger ship is of primary importance. Although diesel engines and propellers are considered to be ones of the main sources of ship noise, our observations have shown that immersed motors contribute significantly to the noise radiated into the water by passenger ships at low speeds. This contribution summarizes a semi-empirical method for estimating the noise radiated into the water from an azimuth thruster-type structure. An analysis of the vibroacoustic behavior in water of the structure excited by an internal acoustic source is detailed. Beyond the complexity of performing full-scale measurements in a noisy environment, the analysis of the measurements allows to estimate the vibration modes of the immersed structure. A numerical model of the dedicated structure, calibrated on the basis of an experimental modal analysis, is used to estimate the pressure field radiated by the structure into the water. As a result of these analyses, a simplified model of radiated noise in the water is developed for use in nominal thruster operation.

Wire-to-two-drop plasma thruster: Experimental and numerical investigation of electroaerodynamic jet flow for micro aerial vehicle propulsion

Conventional micro aerial vehicles (MAVs) have primarily relied on complex, flapping-wing mechanisms for propulsion, often exhibiting limitations in terms of reliability and efficiency. To overcome these challenges, this study explores the potential of electroaerodynamic (EAD) thrusters as a novel propulsion system. By accelerating air molecules through ion collisions, EAD jet flow generates thrust, offering advantages such as noiseless operation and zero emissions due to its moving-part-free design. This research presents a comprehensive experimental and numerical investigation of a wire-to-two-drop thruster configuration to elucidate its electromechanical performance, plasma flow dynamics, and EAD jet characteristics. Experimental measurements of key parameters, including current, thrust, power, and effectiveness, were correlated with numerical simulations, demonstrating excellent agreement with a maximum error below 5%. These findings align strongly with established theoretical frameworks, revealing an inverse square root relationship between effectiveness and thrust. To optimize thruster performance, optimal operating voltages were identified at approximately 8.2, 9.4, and 11.6 kV for inter-electrode gap distances of 10, 15, and 20 mm, respectively, achieving a balanced trade-off between thrust and effectiveness. Detailed numerical visualizations of the plasma flow field, including velocity distribution, jet morphology, potential distribution, and electric field lines, provided valuable insights into the thruster's operation. Building upon these insights, a proof-of-concept EAD flier was constructed and tested, incorporating a serrated emitter electrode and lightweight materials. This flier achieved a mass of 0.5 g and generated a thrust of 0.77 g at 15 kV, resulting in a thrust-to-weight ratio of 1.54 and successful liftoff. This demonstration highlights the potential of EAD propulsion for practical MAV applications.

Analytical model of a Hall thruster

Hall thrusters are one of the most successful and prevalent electric propulsion systems for spacecraft in use today. However, they are also complex devices and their unique E×B configuration makes modeling of the underlying plasma discharge challenging. In this work, a steady-state model of a Hall thruster is developed and a complete analytical solution presented that is shown to be in reasonable agreement with experimental measurements. A characterization of the discharge shows that the peak plasma density and ionization rate nearly coincide and both occur upstream of the peak electric field. The peak locations also shift as the thruster operating conditions are varied. Three key similarity parameters emerge that govern the plasma discharge and which are connected via a thruster current–voltage relation: a normalized discharge current, a normalized discharge voltage, and an amalgamated parameter, α¯, that contains all system geometric and magnetic field information. For a given normalized discharge voltage, the similarity parameter α¯ must lie within a certain range to enable high thruster performance. When applied to a krypton thruster, the model shows that both the propellant mass flow rate and the magnetic field strength must be simultaneously adjusted to achieve similar efficiency to a xenon thruster (for the same thruster geometry, discharge voltage, and power level).

Selection of the method for determination of ignition delay of hypergolic propellants

Ignition delay is one of the most important parameters characterising hypergolic propellants. This parameter has a strong im-pact on thruster operation, especially during the cold start. Ignition delay influences the intensity of pressure rise and its peak values during the start of a thruster. High-pressure levels cause stress inside the chamber wall, which directly affects durability and safety. One of two measurement techniques is usually chosen to determine the ignition delay: visual and pressure-based methods. Visual methods are based on high-speed imaging and subsequent image analysis. In the pressure-based method, the pressure trace is analysed. In this study, both techniques were used together and compared in terms of ignition delay determination of hypergolic propellants igniting during the drop tests. The advantages and disadvantages of both techniques were indicated and described. In the setup used in the study, the visual method was found to be more accurate and reliable.

Deposition characteristics and influence regularity of sputtered products in a low-power RF gridded ion thruster

RF ion thruster has a broad prospect of development via its simple structure, high specific impulse and electrode-less discharge. During the RF operating process, an important but ignored generally issue is that the high-energy metal particles generated by sputtering grid system can diffuse towards the ionization chamber and then deposit on the wall to form a conductive film. As a result, the RF power feed can be significantly inhibited which affects the thruster perform and reliability. In this work, the axial distribution of film thickness on the ionization chamber has been measured which has the characteristic of nonlinear distribution. And the film roughness can be significantly affected due to the gradient of film thickness and deposition temperature. Another corresponding measurement of plasma parameter evolutions in the ionization chamber has been carried out within 6000 min of thruster operation. Results shows that the presence of the deposited film can result in a ∼30 % decrease in the plasma density, which weakens the grid sputtering erosion and subsequently leads to a decrease in film growth rate. The combination of contributed to the degradation of the thruster performance, with the effect of deposited film accounting for about 41 %.

Indicator of abnormal cathode electron emission state with gas flow in Hall thrusters

The accurate diagnosis of abnormal electron emission state in hollow cathodes is crucial for the stable operation of Hall electric propulsion systems. In this study, a method of reducing the cathode working gas flow rate was used to simulate abnormal working conditions in which the cathode electron emission state (CEES) was deteriorating. By analyzing and comparing the oscillation signals under abnormal and steady-state working conditions, it was found that as the CEES deteriorated, the power content of the breathing oscillation decreased in the 1–40 kHz frequency band, and the main frequency decreased; in contrast, the power content of the transit-time oscillation increased in the 100–500 kHz range, and the main frequency was on the rise. Combined with the current growth rate analysis of breathing and transit-time oscillations, when the cathode gas flow rate decreases, the CEES deteriorates, the coupling voltage drop increases, and the potential drop in the channel decreases. The electron temperature and nonlinear power absorption of the electrons decrease, leading to a decrease in the growth rate of breathing oscillations and the breathing oscillation weakens; however, the time-averaged ion velocity and ion sound velocity in the channel decrease simultaneously, but the ion velocity decreases significantly faster than the ion sound velocity, leading to an increase in the growth rate of the transit-time oscillation, and the transit-time oscillation strengthen. Through comparison of the oscillation signals under different working conditions, such as varied anode flow rate, anode voltage, magnetic induction, it was proven to be a unique feature of CEES deteriorates, and can be used as an indicator of CEES deteriorates during the on-orbit operation of the Hall-effect thrusters.

Argon admixture-driven enhanced ionization and performance of a 5 kW Hall thruster on krypton

Utilizing alternative propellants has been recognized as a strategy to reduce the total cost of propellants in electric propulsion-based missions. The aim of this study is to quantify the Hall effect thruster (HET) performance and operating characteristics using a Kr–Ar mixture to enable mission designers to evaluate the impact on mission and spacecraft design. We present the performance and plume plasma properties of the P5 5 kW-class HET operated with a Kr–Ar mixture with Ar volumetric flow rate fractions from 0 to 100%. The thruster is characterized at discharge power levels of 2.6 kW and 4.1 kW at constant discharge current and voltage over the range of Ar fractions. Despite higher ionization energy and lower mass of Ar, the thruster exhibited a similar level of thrust within 2% when comparing the pure Kr and 26%-Ar mixture cases. The derived ion energy distribution functions and analytical modeling suggest that the characteristic length for the ionization region is extended as the Ar fraction increases. The increased residence time of Kr at the extended ionization region and background energetic electrons from the ionized Ar neutrals are considered to cause this enhanced ionization of the injected Kr neutrals. This leads to a 6% higher Kr ion density at the 26%-Ar case even at the 16% less injected Kr neutral density than in the pure Kr case. The enhanced Kr ionization and generated Ar ions in the 26%-Ar case consequently led to a comparable thrust with that of the pure Kr case. The study indicates that mixing Ar with approximately 26% volumetrically with Kr can provide a similar or even higher thrust performance at the same discharge power. This will be particularly advantageous for various space missions that require high impulses by reducing the total cost of the propellant.

Experimental investigation of the polarity-switching process with different bipolar ionic liquid thruster operating frequencies

The bipolar ionic liquid thruster employs ionic liquid as a propellant to discharge positively and negatively charged high-energy particles under an alternating current (AC) power source, effectively suppressing electrochemical reaction and ensuring charge neutrality. Determining an optimal AC supply power source frequency is critical for sustained stable thruster operation. This study focuses on the emission characteristics of the ionic liquid thruster under varied AC conditions. The AC power supply was set within the frequency range of 0.5–64 Hz, with eight specific frequency conditions selected for experimentation. The experimental results indicate that the thruster operates steadily within a voltage range of ±1470 to ±1920 V, with corresponding positive polarity current ranging from 0.41 to 4.91 μA and negative polarity current ranging from −0.49 to −4.10 μA. During voltage polarity switching, an emission delay occurs, manifested as a prominent peak signal caused by circuit capacitance characteristics and a minor peak signal resulting from liquid droplets. Extended emission test was conducted at 16 Hz, demonstrating approximately 1 h and 50 min of consistent emission before intermittent discharge. These findings underscore the favorable impact of AC conditions within the 8–16 Hz range on the self-neutralization capability of the ionic liquid thruster.

Enhancement of the electrostatic–magnetic hybrid thrust performance using a stagnant ring

Electrostatic–magnetic hybrid ion acceleration has been researched to realize high-thrust density and efficient thrust operations. In this study, a newly designed stagnant ring (SR) was installed in an electrostatic–magnetic hybrid thruster with the nominal anode inner radius of 40 mm to improve the thrust performance, and the effects of SR on the thruster operation of xenon propellant were analyzed. During constant-discharge-voltage operations, the discharge current increased with both electrically conductive (copper) and non-conductive (ceramic) SR materials, yet the thrust increased only with the copper SR, which was equipotential to the anode. At lower values of the radial width (H) of the copper SR, the thrust and discharge current were almost dominated by the propellant flow rate from the anode surface. However, the contribution of the propellant flow rate from the cathode tip on the thrust and discharge current became significant when H ≥ 9 mm. The thrust coefficient in electromagnetic acceleration form exhibited a maximum value at H = 9 mm, whereas that in electrostatic acceleration form monotonically increased when H ranged from 0 to 20 mm. The highest thrust efficiency was obtained at H = 12 mm. The study findings indicate that installing a conductive SR can effectively enhance the thrust performance in electrostatic–magnetic hybrid thrusters.

Non-intrusive measurements of plasma impedance in an electron-cyclotron resonance thruster

We present a general method allowing for direct measurements of the complex impedance of radio frequency devices that need to be powered by an external power supply unit, making use of a dual-directional coupler and a vector network analyzer. This method is used to measure the impedance of an operating electron-cyclotron resonance thruster, which is then calibrated using a custom triple-short calibration kit to retrieve the impedance of the plasma within the chamber. The measurements, realized without disturbing the thruster operation, show that the plasma impedance ranges between 55 and 90 Ω in resistance and between −30 and 20 Ω in reactance. An equivalent circuit model is finally discussed, aiming at estimating plasma parameters through these impedance measurements.

A review of the impact of ground test-related facility effects on gridded ion thruster operation and performance

A key consideration in the interpretation of ground test data of electric propulsion devices purposed for spaceflight is understanding how facility-effects influence thruster operation. This understanding is critical to the prediction of actual thruster performance in space. The necessity of science-based predictions gleaned from ground tests are particularly critical at higher thruster power levels. Operation of engines at higher power levels in vacuum chambers leads to considerable elevation in background pressure, background plasma density, and backsputter rates. This review examines the influence of ground test facility effects on gridded ion thruster operation. Ground test operation is compared with flight data, where available, to obtain a clear picture of operational differences. Mitigation strategies to alleviate facility effects are also commented upon.

Research on the hydrodynamic performance of propellers under oblique flow conditions

The phenomenon of oblique water inflow is widespread in the operation of thrusters, which will cause adverse effects on the hydrodynamic performance of thrusters. In order to investigate the hydrodynamic performance and stress variation of the propeller in oblique flow, based on computational fluid dynamics theory, this paper takes the DTMB4119 propeller as the research object and conducts numerical simulation research on the propeller in oblique flow by solving the RANS equation. By calculating the open water performance curve and surface pressure distribution of the propeller, the rationality of the numerical method and the meshing are verified. Considering the flow field information such as velocity, oblique flow angles, flow line distribution, and pressure distribution, the changes in hydrodynamic characteristics of the propeller are simulated and analyzed. The results show that the uneven distribution of pressure on the propeller blade surface increases with a decrease in the advance coefficient. The force pulsation amplitude of a single blade increases with an increase in oblique flow angles. With the increase in oblique flow angles, the thrust, torque, and efficiency of the propeller show different increasing trends. With the increase in propeller advance coefficient, propeller thrust and torque decrease gradually, and propeller efficiency increases gradually. By using mature propellers, this paper has more reference value for studying the flow field around the ship hull and the hydrodynamic performance of propellers in the process of ship maneuvering.

Fully discrete model of kinetic ion-induced electron emission from metal surfaces

Ion-induced electron emission (IIEE) is an important process whereby ions impinging on a material surface lead to net emission of electrons into the vacuum. While relevant for multiple applications, IIEE is a critical process of electric thruster (ET) operation and testing for space propulsion, and, as such, it must be carefully quantified for safe and reliable ET performance. IIEE is a complex physical phenomenon, which involves a number of ion-material and ion-electron processes, and is a complex function of ion mass, energy, and angle, as well as host material properties, such as mass and electronic structure. In this paper, we develop a discrete model of kinetic IIEE to gain a more accurate picture of the electric thruster chamber and facility material degradation processes. The model is based on three main developments: (i) the use of modern electronic and nuclear stopping databases, (ii) the use of the stopping and range of ions in matter to track all ion and recoil trajectories inside the target material, and (iii) the use of a scattering Monte Carlo approach to track the trajectories of all mobilized electrons from the point of first energy transfer until full thermalization or escape. This represents a substantial advantage in terms of physical accuracy over existing semi-analytical models commonly used to calculate kinetic IIEE. We apply the model to Ar, Kr, and Xe irradiation of W and Fe surfaces and calculate excitation spectra as a function of ion depth, energy, and angle of incidence. We also obtain minimum threshold ion energies for net nonzero yield for each ion species in both Fe and W and calculate full IIEE yields as a function of ion energy and incidence angle. Our results can be used to assess the effect of kinetic electron emission in models of full ET facility testing and operation.

Current Sheet Evolution in a Planar Inductive Pulsed Plasma Thruster

Abstract The evolution of the current sheet is fundamental to the understanding and operation of planar inductive pulsed plasma thrusters. Methods for experimentally determining the time-varying mutual inductance and plasma resistance associated with the current sheet are presented. From these, time-histories of the resistive heating in and electromagnetic acceleration of the sheet are found. Analysis of experimental data obtained from a compact, low discharge energy inductive pulsed plasma thruster shows that current sheet evolution is well described by three phases: ionization, formation, and acceleration. Evidence is provided for the plasma current and resistive heating becoming localized near the upstream edge of the current sheet as the sheet forms. The influence of initial propellant density and pre-ionization on the characteristics of the sheet are examined. Both higher propellant densities and increased pre-ionization are found to produce current sheets which exhibit greater magnetic impermeability. Higher densities cause the sheet to form closer to the coil face, improving coupling, while stronger pre-ionization leads to more rapid resistive heating, causing the sheet to form earlier in time.

Xenon tracers for cost effective laser induced fluorescence of alternative propellant Hall thrusters

One of the limiting factors to developing plasma thrusters on alternative propellants is the cost associated with changing the diagnostic tools, which are often propellant-dependent. For laser induced fluorescence (LIF), which is typically used for ion velocity distribution measurements to determine ion trajectories and potential profiles, either new lasers need to be bought, which are tuned to the wavelength of the new element’s excitation level, or a costly tunable laser is required. A method to use existing LIF setups designed for xenon on any propellant has been demonstrated on a Hall thruster operating on krypton. In the demonstration test, a small amount of xenon (0.01%–4%) was mixed with the main krypton propellant for use as a diagnostic tracer, and xenon ion velocities were measured while also monitoring changes in the mean discharge current and oscillations. High signal-to-noise ratios in LIF data acquired along the channel centerline were obtained with tracer gas fractions ≤1% that negligibly affected the thruster operation. These results and comparison of the emission spectra of xenon and other common propellants suggest that the tracer LIF method should be broadly applicable to LIF measurements in Hall thrusters operating on alternative propellants.

Implementation of direct battery driven discharge in triggerless operation of vacuum arc thrusters

The increasing demand for CubeSat missions with limited budgets and constrained timelines necessitates the development of innovative propulsion systems. This paper focuses on vacuum arc thrusters (VATs) and investigates the implementation of battery-driven discharge for triggerless firing. VATs offer a promising alternative for CubeSat propulsion, generating thrust through the pulsed ejection of plasma in a cathode arc discharge. The study explores the feasibility of direct battery-driven discharge in VATs and demonstrates stable and variable-duration pulses. Experimental testing using high C-rating lithium-ion (Li-ion) batteries showcase the advantages and operational benefits of this approach. The results highlight the potential for variable energy pulses, precise control of thrust impulse bit, and improved performance compared to capacitor-driven systems. Future research directions include optimising pulse duration for enhanced lifetime, plume analysis, and erosion rate investigations. The implementation of VATs with battery-driven discharge presents a promising solution to enhance the performance and reliability of CubeSat propulsion systems while ensuring compliance with space sustainability guidelines.

Experimental investigation into efficiency loss in rotating magnetic field thrusters

An experimental investigation into the low thrust efficiency of a rotating magnetic field (RMF) thruster is presented. This technology is a low maturity but potentially enabling candidate for high-power in-space propulsion for use with alternative propellants. Direct thrust stand measurements of a 5 kW class RMF thruster were performed and show the thrust efficiency was 0.41 ± 0.04% with a specific impulse of 292 ± 11 s—typical values for RMF thruster operation. A suite of far-field probes were used to inform a phenomenological efficiency model for RMF thruster performance that accounted for divergence, power coupling, mass utilization, and plasma/acceleration efficiency. It was found that the plasma efficiency was the critically low term at 6.4 ± 1.0%. This indicates that the majority of the energy coupled to the plasma from the RMF antennas was lost before being converted to directed kinetic energy in the thruster beam. To determine the source of these losses, time-resolved measurements of the internal plasma properties were performed using a triple Langmuir probe. It was found that collisional excitation radiation and wall losses were the two dominant loss processes. This trend can be explained by the unusually high plasma density ( >1019 m−3) exhibited by this device compared to other electric propulsion architectures. Limitations in the probing techniques and strategies for improving RMF thruster performance are discussed given the results from the efficiency analysis.

Effects of propellant species on the discharge characteristics of glow discharge hollow cathode

Hollow cathodes are the key sources of electrons for the operation of space electric thrusters. Traditionally, hollow cathodes utilize heating elements to raise the temperature of low work function materials to produce electrons. To accommodate the development of small satellites, researchers around the world have developed alternative neutralizer technologies such as radio frequency neutralizers, electron cyclotron resonance neutralizers, and coiled tungsten filaments cathode neutralizers to simplify the cathodes design and operation. In this work, a lower current glow discharge hollow cathode (GDHC) is developed, which can generate a discharge current of 10–100 mA. By testing with xenon, krypton, and argon propellants, it is found that GDHC has lower discharge voltage and power consumption with argon propellant. Simulation analysis reveals that during argon propellant discharge, the secondary electron emission coefficient on the cathode surface is larger, the electron density in the cathode sheath area is higher, and the high mobility of argon ions increases the conductivity within the cathode sheath. In addition, there are low and high extraction modes for GDHC electron extraction, depending on whether the keeper electrode absorbs or emits electrons.

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 (<0.1 A) to higher-current (>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.