Thruster Parameters Research Articles

Analysis and control of Hall effect thruster using optical emission spectroscopy and artificial neural network

This study presents a proof-of-principle for using optical emission spectroscopy and artificial neural networks for real-time monitoring and control of the operational parameters of a Hall effect thruster: the anode voltage, the anode xenon injection, the discharge current, and the coil current. In that regard, we build an optical database of 26 spectral lines across 6469 operating conditions to train and test the neural network. We then reduced the learning lines from 26 to 15 based on their statistical correlation with the target parameters. After tuning the hyperparameters of the network, the network predicted the thruster’s parameters with notable accuracies: 95% for the anode voltage, 84% for the coil current, and 99% for both the anode flow rate and the discharge current. The estimated uncertainty of predictions, at 3σ, is ±51V for voltage, ±1A for coil current, ±0.15A for discharge current, and ±0.15mgs−1 for anode flow rate. The prediction calculations were within milliseconds and enabled real-time monitoring of the thruster parameters. Therefore, a proportional-integrator-derivative controller (PID) controller was implemented to regulate the anode voltage and flow rate based on the optical emission of the plume. The PID showcased short settling times from 0.1 to 0.4 s and overshoot levels up to 3% of the target value for the voltage and 10% of the target value for the flow rate. These results were for a fixed coil current at 4A. The study showed that changing the coil current may necessitate more sophisticated prediction models and control strategies. Future work will expand the model’s generalizability to different thruster types, propellants, and magnetic field configurations.

Fault Tolerant Control in Underwater Vehicles

The article discusses the problem of adaptive system synthesis to provide the detection, estimation, and elimination of the negative consequences of faults in the thrusters of underwater robots and deviations of their parameters. For this purpose, an approach consisting of three stages is proposed. At the first stage, based on a bank of diagnostic observers, deviations of the thruster parameters from their nominal values or the occurrence of errors in the readings of its sensors are detected. At the second stage, the magnitudes of the detected deviations and errors are estimated using sliding mode observers. To estimate not only single but also multiple faults, the additional sliding mode observers are used. At the third stage, a control signal is generated to retain the basic dynamic properties of the thruster based on the estimated magnitudes of deviations and errors. Thus, the main contribution of this paper is designing the systems compensating for the consequences of faults occurring in the thrusters of the UR and errors in the readings of the angular velocity sensors. This avoids reducing the accuracy of movement along specified trajectories, as well as preventing accidents that could lead to the loss of the vehicle. The operability and efficiency of the proposed systems were tested using computer simulations and experimental studies on an electro-mechanical stand.

Optimization of the accumulator tank filling modes of the xenon feed system for electric propulsion system

The object of the study in this work is pressure spikes in the accumulator tank of the xenon feed system for electric propulsion. The presence of significant pressure spikes in the accumulator tank of the feed system causes a change in the mass flow and, accordingly, a significant deterioration of the thruster parameters - thrust, specific impulse and efficiency. The problem that was solved in this work is to find ways to reduce pressure spikes while minimizing the volume of the accumulator tank without using additional elements. The analysis of literary sources showed that the specified problem is typical and has not been solved for small-volume accumulator tanks yet. To solve this problem, theoretical and experimental determination of pressure spikes that occur during the work on feedback from sensors was carried out, and an improved method of filling the accumulator tank was proposed. As a result of research, it was determined that pressure spikes that go beyond the permissible range (±3 %) appear at an inlet pressure of 3.5 MPa. With an increase of inlet pressure, the pressure spikes also increase. As a result of research, it was found that for an inlet pressure of up to 3 MPa, it is appropriate to work on feedback from pressure sensors. For pressure values from 3 to 6 MPa, it is necessary to use the experimentally obtained formula to determine the accumulator tank filling time. For a pressure of more than 6 MPa, it is necessary to work in the bang-bang mode. The research made it possible to determine the optimal modes of filling the accumulator tank. Presented modes allow to use a small-volume accumulator tank with the minimization of the valve activations number, in which stable operation of the thruster is maintained. The conclusions obtained as a result of the research can be useful for most developers of feed systems for electric propulsion.

ANALYSIS OF THE EFFECT OF CHANGING THE WORKING SUBSTANCE FROM XENON TO ALTERNATIVE INERT GASES ON THE PARAMETERS OF THE HALL-TYPE ELECTRIC PROPULSION SYSTEM

Xenon is the main working substance for Hall-type electric propulsion systems (EPS) because it allows to obtain high performance of the EPS with ease of operation. However, the high cost of xenon affects the overall cost of development and operation of the EPS system with xenon. Therefore, research is being conducted worldwide to explore alternative working substances that can reduce the cost of tests and development of the EPS while maintaining acceptable performance. Transitioning from xenon to an alternative working substance will lead to changes in the thruster parameters, and as a result, requires changes and modifications of EPS systems for exact working substance. This study analyzes the impact of transitioning from xenon to alternative inert gases such as krypton and argon on the Hall-type EPS, including the thruster, power processing unit, working substance storage andfeed system. It is established that the transition to alternative inert gases, even with thruster optimization for the new working substance, leads to a deterioration of all thrusterparameters except specific impulse.The increased power consumption of the thruster will require the development of a new power processing unit. However, if the thruster operates on an alternative working substance with the same power as with xenon, the power processing unit may remain unchanged. The lower density of alternative working substances would necessitate an increase in the storage tank volume, and the lower critical temperature would eliminate the need for heaters to prevent the ingress of the working substance's liquid phase into the accumulator tank of the feed system.In the feed system, only flow restrictors will need to be replaced to ensure the required flow rate of the alternative working substance. It is found that the minimal modifications required for adapting the EPS to an alternative working substance isoccurred when the thruster operates with the same power consumption as it does with xenon.

PARAMETERS OF THE ST-40M HALL THRUSTER WITH INCREASED POWER DISCHARGE SUPPLY


 
 
 Abstract. The article presents the results of experimental investigation of the ST-40M Hall thruster modification developed by the Space Electric Thruster Systems (SETS). The thruster used a new filament hollow cathode with an emitting insert made of lanthanum hexaboride (LaB6). The use of a new hollow cathode made it possible to study the thruster’s parameters in the discharge power range of 300 – 550 W. The investigation was carried out using a stabilized voltage supply as a discharge power supply. As a result of experimental investigation, the current-voltage characteristics of the discharge in the accelerating channel of the thruster were obtained at fixed values of the anode’s mass flow rate of the working gas - xenon. The experimental dependences of thrust on the anode’s mass flow rate of the working substance and the discharge voltage, as well as the dependences of the specific impulse and thruster efficiency on the discharge voltage, are obtained. Experimental studies have shown that with a change in the discharge voltage in the range of 250 – 400 V and a change in the mass flow rate of the working gas in the range of 0.9 – 2.2 mg/s, the thruster provides a thrust value of 12 – 33 mN, a specific impulse value of 1200 – 1800 s, and the maximum value of the anode’s efficiency is 48 – 52%. The article also contains oscillogramsillustratingtherepeatedprocessofthethrusterstarting. Experimentalinvestigationofthemodified thruster ST-40M with the new hexaboride hollow cathode showed a significant improvement in thruster parameters compared to the prototype parameter before modification. In the process of experimental investigation, it was found that in some modes of the thruster operation, significant fluctuations in the discharge current occur, which led to a significant decreasing in the thruster parameters. For a more accurate clarification of the nature and parameters of the discharge current oscillations, it is necessary to conduct additional studies. The obtained experimental parameters of the modified ST-40M thruster made it possible to determine the values of its optimal parameters for various space missions.
 
 

A comparative study on the different energy supply modes for a micro-cathodic arc thruster

The energy supply mode of a micro-cathodic arc thruster (μ-CAT) was adjusted based on the charge and discharge characteristics of the energy storage element, and the cathode discharge characteristics of μ-CAT were studied under different energy supply modes. We applied an electrical method to measure the discharge characteristics of μ-CAT, with an enhanced intensified charge coupled device spectrum detector to measure the plasma parameters of μ-CAT, allowing analysis of the internal logic among thruster plasma parameters, arc discharge characteristics, and thruster energy supply mode. The mechanisms of μ-CAT were analyzed and characteristics of accumulated discharge, energy distribution, and ion velocity distribution of thruster were compared under different energy supply modes to obtain the effect of the energy supply mode on the discharge characteristics of the interelectrode of μ-CAT. The results show that the energy supply mode of μ-CAT affected the ionization characteristics and ionization stability of the micro-cathodic arc thruster, causing a multistage ionization difference in the cathode metal target. This would not only affect the propulsion performance of the thruster, but also affect the change of the number of particles in the cathode region in a single discharge cycle.

Detection of the magnetic field parameters influence on discharge current fluctuations and optimal operation modes of the hall thruster

The object of research reported in this work is the fluctuations of the discharge current in Hall thrusters. The presence of significant fluctuations in the discharge current in Hall thrusters causes a significant deterioration of the thruster parameters – thrust, specific impulse, and efficiency. The task addressed in the current work relates to determining the main factors that affect the conditions for the occurrence of discharge current fluctuations, finding ways to reduce them, and obtaining optimal values of the parameters of the Hall thruster. The review of literary sources revealed that the specified problem is typical and has not yet been solved. In order to solve this problem, studies of the ST-40M Hall thruster were carried out in order to determine the main factors that have the greatest influence on the occurrence of oscillations and means of reducing the oscillations of the discharge current. The result of the research determined that the fluctuations of the discharge current depend most significantly on the parameters of the magnetic field in the acceleration channel of the thruster. The parameters of the magnetic field are determined by the magnitude of the currents of the thruster electromagnet coils, and the nature of oscillations, amplitude, and frequency may differ significantly with a slight change in the coil currents. As a result of the study, it was found that for the values of the currents of the coils of the magnetic system of the thruster, for which the level of fluctuations of the discharge current is minimal, the average value of the discharge current of the thruster also takes a minimum value. The research made it possible to determine the optimal operating modes of the Hall thruster, which ensure, at a given level of specific power, the maximum values of thrust, specific impulse, and thruster efficiency. The conclusions from the study could be useful for most developers of electric propulsion systems based on Hall thrusters

On Propulsion-based Estimation of Incoming Flow Velocity

AUV’s actuators systems are highly affected by external conditions such as incoming flow velocity. The paper aims to estimation of incoming flow velocity based on quasi-steady model of thruster. The work motivation is compensation of incoming flow effect on AUV’s actuators dynamic model for improving actuator control or a velocity estimation in case of velocity sensors absence. The equation for incoming flow velocity estimation is presented. It accounts orientation of thrusters relative to the longitudinal vehicle axis. Thruster parameters identification procedure based on bollard pull tests and marine trials are presented. Comparative results of propulsion-based velocity estimation with DVL data are discussed. Velocity estimation error between methods is less than 0.1 m/s in steady motion and less than 0.2 m/s in transition process.

Faraday cup sizing for electric propulsion ion beam study: Case of a field-emission-electric propulsion thruster.

This article provides information about the sizing and standardization of a Faraday cup (FC) used as a plasma diagnostic. This instrument is used to accurately map the ion beam profile produced by an electric propulsion (EP) device. A FC is a cylindrical probe that uses an electrode, termed collector, to measure the current. Several studies have shown the relevance of adding an extra electrode, called collimator, to define the collection area and to minimize interactions with the ambient plasma. Both the electrodes are encapsulated into an isolated metallic housing that prevents ambient plasma from disturbing the measurements. In this case study, a field-emission-electric propulsion (FEEP) thruster is used. The FEEP technology uses electrostatic fields to extract liquid metal (indium) ions from a sharp surface and accelerates them to high velocities, providing thrust. The FEEP model used in this study is the ENPULSION NANO thruster from the Austrian company Enpulsion. We present results focusing on the sizing of a FC in terms of cup length, aperture diameter, and collection solid angle as well as on the material exposure to the ion beam. For a far-field ion beam study of a FEEP indium based electric thruster, our study outcomes show that a FC optimum sizing is a 50mm long collector cup and a 7mm wide inlet aperture. Moreover, shielding the repeller/collimator from direct exposure to the ion beam seems to greatly minimize perturbation during ion current acquisition. Finally, to only measure the ion current, a negative potential should be applied to the collector and repeller, where the latter is more negative. This study contributes to the effort on diagnostic standardization for EP device characterization. The goal is to enable repetitive and reliable determination of thruster parameters and performances.

IFM Nano Thruster performance studied by experiments and numerical simulations

Field emission electric propulsion thrusters are characterised by their low thrust range, which makes them ideal for the precise control of spacecraft. Decisive for precise control are the properties of the thruster ion beam, which includes the beam divergence angle and the thrust vector. The analysis of these properties is also necessary in order to be able to estimate the interactions of the beam with components of the spacecraft. Due to such interactions, solar panels or electrical instruments on board the spacecraft could be damaged by sputtering effects.The spatial ion current density and energy distribution of a test crown emitter beam, with different specifications compared to the IFM Nano thruster, were examined experimentally with a diagnostic system, including Faraday cups and a retarding potential analyser. In addition to the analysis of the beam profile of an emitting crown, a single emitting needle was analysed. Based on these experimental analyses, an ion trajectory simulation model was developed to determine the theoretical ion current density distribution. This model includes the properties of a liquid metal ion source, where the ion trajectories start from their point of origin, the so-called Taylor cone jet cap. The benchmark of the model shows that the thrust vector and divergence angle correspond to the experimental results and shows the identical calculations for different thruster parameters, like emission current and electrode voltage. The simulation allows for the optimisation of existing and novel thruster geometries in terms of performance, reliability and longevity.

Particle-in-cell numerical simulation of the PHall-IIc Hall thruster

In this work we compare the performance of two types of Hall thrusters, namely, the SPT-100 and the PHall-IIc, the latter being developed by the Plasma Physics Laboratory at the University of Brasilia. The comparison is carried out by performing particle-in-cell simulations. We compute thruster parameters such as thrust and specific impulse. Our results can contribute to the design of electric propulsion devices for future Brazilian space missions.

Optimization of discharge parameters in an inductive RF ion thruster prototype

This paper presents the results of an experimental study of inductive RF ion thruster prototype parameters, which have demonstrated a strong dependence of the prototype characteristics on the length of the gas-discharge chamber and the operational frequency of the RF generator. The best result was obtained when the prototype was operated at a frequency of 4 MHz with a gas-discharge chamber 5 cm long. The experiments demonstrated that there is a significant resource for improving thruster parameters associated with an increase in the fraction of the RF power absorbed by the plasma. The latter can be achieved by superimposing an external magnetic field on the discharge. In parallel with the results of the experiments, this paper presents numerical calculations based on the model taking into account the presence of a parasitic capacitive channel of the RF power input into the discharge, as well as power losses in the external circuit. Calculations have shown that the experimentally observed effects are associated with the effect of the capacitive RF power input channel, as well as the frequency dependence of the plasma's ability to absorb the RF power.

Optimization, Test and Diagnostics of Miniaturized Hall Thrusters.

Miniaturized spacecraft and satellites require smart, highly efficient and durable low-thrust thrusters, capable of extended, reliable operation without attendance and adjustment. Thermochemical thrusters which utilize thermodynamic properties of gases as a means of acceleration have physical limitations on their exhaust gas velocity, resulting in low efficiency. Moreover, these engines demonstrate extremely low efficiency at small thrusts and may be unsuitable for continuously operating systems which provide real-time adaptive control of the spacecraft orientation, velocity and position. In contrast, electric propulsion systems which use electromagnetic fields to accelerate ionized gases (i.e., plasmas) do not have any physical limitation in terms of exhaust velocity, allowing virtually any mass efficiency and specific impulse. Low-thrust Hall thrusters have a lifetime of several thousand hours. Their discharge voltage ranges between 100 and 300 V, operating at a nominal power of <1 kW. They vary from 20 to 100 mm in size. Large Hall thrusters can provide fractions of millinewton of thrust. Over the past few decades, there has been an increasing interest in small mass, low power, and high efficiency propulsion systems to drive satellites of 50-200 kg. In this work, we will demonstrate how to build, test, and optimize a small (30 mm) Hall thruster capable of propelling a small satellite weighing about 50 kg. We will show the thruster operating in a large space environment simulator, and describe how thrust is measured and electric parameters, including plasma characteristics, are collected and processed to assess key thruster parameters. We will also demonstrate how the thruster is optimized to make it one of the most efficient small thrusters ever built. We will also address challenges and opportunities presented by new thruster materials.

A Low-Cost Test Bench for Underwater Thruster Identification

Underwater vehicles are increasingly being used for the inspection of marine structures, bathymetry and underwater intervention applications. The manoeuvring performance of these vehicles is critically dependant on the quality of the thruster model which must faithfully represent the relationship between the thruster inputs and corresponding outputs. Thruster manufacturers typically provide only the essential electrical, mechanical and thrust performance characteristics, for this reason the identification of the thruster’s parameters is especially important for constructing a reliable control system for the underwater vehicle. Physical testing using some form of test bench is normally used for identifying the thruster parameters. These test benches usually consist of a water tank and the thruster which is clamped to some form of instrumented reaction frame for measuring the model parameters. The cost of a test bench varies according to two major factors which are the size and the instrumentation which strongly depends on the physical quantities that we are trying to measure. In this article we will make a state of the art survey of existing test benches in order to provide a technical basis that will allow us to build a low-cost test bench. A thruster will be modelled using the resulting test bench and the outcomes will be presented at the end of this paper.

Particle-in-cell simulations of Hall thruster acceleration and near plume regions

The results of 1-D azimuthal, and 2-D z-θ, particle-in-cell simulations of Hall thruster acceleration and near plume regions are reported. The 1-D simulations show that fast moving electron waves drive coherent, large amplitude, ion acoustic waves. The ion waves become azimuthally coherent too quickly for ion sound to be the cause; the coherence comes from early electron waves. 2-D simulations with only fast moving, main beam, ions result in ion acoustic wave amplitudes much lower than those in the 1-D calculations. The addition of a small amount of ionization in the domain, at a rate consistent with Hall thruster parameters, causes the potential profile to divide into two distinct regions. In the upstream portion of the domain, the potential profile is very steep; in the downstream portion, the potential profile is almost flat. Electron transport in the two regions is discussed.

On the Effect of the Surface State and Geometry of the Discharge- Chamber Edges and Sizes of Morozov’s Stationary Plasma Thruster upon its Long-Term Operation

The results of long-term tests of Morozov’s stationary plasma thrusters are presented. It is revealed how the surface state and geometry of the discharge chamber’s edges influence the thruster’s parameters. It is shown that, during the ground tests of thrusters with cylindrical geometry of the acceleration channel under initial stage of operation, material sputtered from the discharge chambers’ walls is deposited onto the nearanode segment of the walls. Films of deposited material fail during thruster operation causing fragment formation, which jut out towards the discharge volume and disturb the motion of drifting electrons in the area of their acceleration. As a result the thruster reactive force and specific impulse decrease. The way in which the forming fragments influence thruster performance and operation is examined. It is shown that it decreases under long-term operation and significant channel widening since the ion flux to the wall and the quantity of the sputtered material decrease, and since the profile of the walls changes due to their wear and cleaning effect of the discharge. As a result the thruster’s parameters are restored to a level close to the initial one. It is shown that the dynamics of thruster parameters variation in space and during ground tests is different. This means that it is necessary to simulate more properly the conditions of thruster operation in space when conducting ground development tests. Thrusters with a long lifetime should be designed with widening of the acceleration channel beyond the loop which surrounds the magnetic system so that areas of acceleration and the erosion of walls are located in the widened part of the channel.

On the use of time-domain simulation in the design of Remotely Operated Submergible Vehicles

Designing a Remotely Operated Vehicle (ROV) is a complicated task in which the design team deals with a considerable amount of uncertainty before the device is able to be tested at full scale. A way to cope with such uncertainty is to use simulation software to evaluate design concepts along the different levels of abstraction of the process. In this work, the use of aNySIM, the Maritime Research Institute Netherlands (MARIN) multibody time-domain simulation tool, as a part of the design process of an ROV is addressed. The simulation software is able to solve the equations of motion of the vehicle based on rigid body dynamics, including features such as hydrodynamics, hydrostatics, thrusters, thrust allocation, and PID control. Different simulation scenarios are proposed to evaluate different concept solutions to the design, including thruster parameters and distribution. The results are further used to select the concept solutions to be implemented in the final design.

Multiphysics Model for Radio-Frequency Gridded Ion Thruster Performance

A multiphysics radio-frequency gridded-ion thruster performance model is presented. The model is composed of various submodels to account for different physics phenomena. A two-dimensional axisymmetric ion extraction submodel is used to estimate the ion optics effective transparency and beam divergence. A two-dimensional axisymmetric and three-dimensional molecular neutral gas submodels are used to calculate the ion optics Clausing factor and the discharge chamber pressure, respectively. An electron transport submodel determines the rate coefficients and the effective collision frequency by solving a two-term approximation Boltzmann equation. The plasma properties are calculated by employing a volume-averaged zero-dimensional plasma submodel. A fully three-dimensional thruster geometry is used for determining electromagnetic thruster and plasma parameters, such as the coupling coefficient, power losses, and field distributions. The submodel solves Maxwell’s equations in the thruster and plasma represented by a complex conductivity. To improve on the uniform plasma assumption, a radial variation in plasma density/conductivity is introduced. The impedance matching and circuit power losses are accounted for by a radio-frequency circuit submodel, in which the thruster is represented by a complex impedance obtained from the electromagnetic submodel. It is shown that the model predicts the Radio-Frequency Ion Thruster 3.5 experimentally measured radio-frequency generator power and current data values (and trends) within 13% error bound for various operational conditions.

Experimental research on low-cost cold gas propulsion for a space robot platform

The attitude and orbit control system of satellites and space probes plays a crucial role in missions. One commonly used attitude control method relies on small rocket thrusters. This paper focuses on experimental research into a low thrust, cold gas satellite thruster, which is the simplest solution in this group. A dedicated research stand was designed and built to measure the key parameters of the thruster: thrust and mass flow rate. The measurements were used to calculate specific impulse and to compare it against expected values. Dynamic parameters were also identified – delay time and valve opening time, and power consumption of the coil. Minimum impulse bit and maximum frequency of operation were determined through research with pulse width modulation. A multicycle experiment was conducted to investigate the effect of the number of cycles on thruster parameters. A detailed description of the research stand and measurement methods is given, followed by the results.

Perch Landing Assisted by Thruster (PLAT): Concept and Trajectory Optimization

A concept of the perch landing assisted by thruster (PLAT) for a fixed wind aircraft is proposed in this paper. The proposed concept is applicable to relatively large unmanned aerial vehicles (UAV), hence can overcome the limitation of existing perch landing technologies. A planar rigid body motion of an aircraft with aerodynamic and thruster forces and moments is modeled. An optimal control problem to minimize the fuel consumption by determining the histories of thruster and elevator deflection angle with specified terminal landing condition is formulated and solved. A parametric study for various initial conditions and thruster parameters is conducted to demonstrate the practicability of the proposed concept.