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At least 253 records · Page 14
All sourcesNASA NTRSDOE OSTI
NASA NTRS2012

Optical Diagnostic Characterization of High-Power Hall Thruster Wear and Operation

Optical emission spectroscopy is employed to correlate BN insulator erosion with high-power Hall thruster operation. Specifically, actinometry leveraging excited xenon states is used to normalize the emission spectra of ground state boron as a function of thruster operating condition. Trends in the strength of the boron signal are correlated with thruster power, discharge voltage, and discharge current. In addition, the technique is demonstrated on metallic coupons embedded in the walls of the HiVHAc EM thruster. The OES technique captured the overall trend in the erosion of the coupons which boosts credibility in the method since there are no data to which to calibrate the erosion rates of high-power Hall thrusters. The boron signals are shown to trend linearly with discharge voltage for a fixed discharge current as expected. However, the boron signals of the higher-power NASA 300M and NASA 457Mv2 trend with discharge current and show an unexpectedly weak to inverse dependence on discharge voltage. Electron temperatures measured optically in the near-field plume of the thruster agree well with Langmuir probe data. However, the optical technique used to determine Te showed unacceptable sensitivity to the emission intensities. Near-field, single-frequency imaging of the xenon neutrals is also presented as a function of operating condition for the NASA 457 Mv2.

Williams, George J., Jr.↗
NASA NTRS2013

Mode Transitions in Hall Effect Thrusters

Mode transitions have been commonly observed in Hall Effect Thruster (HET) operation where a small change in a thruster operating parameter such as discharge voltage, magnetic field or mass flow rate causes the thruster discharge current mean value and oscillation amplitude to increase significantly. Mode transitions in a 6-kW-class HET called the H6 are induced by varying the magnetic field intensity while holding all other operating parameters constant and measurements are acquired with ion saturation probes and ultra-fast imaging. Global and local oscillation modes are identified. In the global mode, the entire discharge channel oscillates in unison and azimuthal perturbations (spokes) are either absent or negligible. Downstream azimuthally spaced probes show no signal delay between each other and are very well correlated to the discharge current signal. In the local mode, signals from the azimuthally spaced probes exhibit a clear delay indicating the passage of "spokes" and are not well correlated to the discharge current. These spokes are localized oscillations propagating in the ExB direction that are typically 10-20% of the mean value. In contrast, the oscillations in the global mode can be 100% of the mean value. The transition between global and local modes occurs at higher relative magnetic field strengths for higher mass flow rates or higher discharge voltages. The thrust is constant through mode transition but the thrust-to-power decreased by 25% due to increasing discharge current. The plume shows significant differences between modes with the global mode significantly brighter in the channel and the near-field plasma plume as well as exhibiting a luminous spike on thruster centerline. Mode transitions provide valuable insight to thruster operation and suggest improved methods for thruster performance characterization.

interplanetary exploration↗
NASA NTRS2009

Hall-Effect Thruster Simulations with 2-D Electron Transport and Hydrodynamic Ions

A computational approach that has been used extensively in the last two decades for Hall thruster simulations is to solve a diffusion equation and energy conservation law for the electrons in a direction that is perpendicular to the magnetic field, and use discrete-particle methods for the heavy species. This "hybrid" approach has allowed for the capture of bulk plasma phenomena inside these thrusters within reasonable computational times. Regions of the thruster with complex magnetic field arrangements (such as those near eroded walls and magnets) and/or reduced Hall parameter (such as those near the anode and the cathode plume) challenge the validity of the quasi-one-dimensional assumption for the electrons. This paper reports on the development of a computer code that solves numerically the 2-D axisymmetric vector form of Ohm's law, with no assumptions regarding the rate of electron transport in the parallel and perpendicular directions. The numerical challenges related to the large disparity of the transport coefficients in the two directions are met by solving the equations in a computational mesh that is aligned with the magnetic field. The fully-2D approach allows for a large physical domain that extends more than five times the thruster channel length in the axial direction, and encompasses the cathode boundary. Ions are treated as an isothermal, cold (relative to the electrons) fluid, accounting for charge-exchange and multiple-ionization collisions in the momentum equations. A first series of simulations of two Hall thrusters, namely the BPT-4000 and a 6-kW laboratory thruster, quantifies the significance of ion diffusion in the anode region and the importance of the extended physical domain on studies related to the impact of the transport coefficients on the electron flow field.

plasma↗
NASA NTRS2012

Design of a Laboratory Hall Thruster with Magnetically Shielded Channel Walls, Phase III: Comparison of Theory with Experiment

A proof-of-principle effort to demonstrate a technique by which erosion of the acceleration channel in Hall thrusters of the magnetic-layer type can be eliminated has been completed. The first principles of the technique, now known as "magnetic shielding," were derived based on the findings of numerical simulations in 2-D axisymmetric geometry. The simulations, in turn, guided the modification of an existing 6-kW laboratory Hall thruster. This magnetically shielded (MS) thruster was then built and tested. Because neither theory nor experiment alone can validate fully the first principles of the technique, the objective of the 2-yr effort was twofold: (1) to demonstrate in the laboratory that the erosion rates can be reduced by >order of magnitude, and (2) to demonstrate that the near-wall plasma properties can be altered according to the theoretical predictions. This paper concludes the demonstration of magnetic shielding by reporting on a wide range of comparisons between results from numerical simulations and laboratory diagnostics. Collectively, we find that the comparisons validate the theory. Near the walls of the MS thruster, theory and experiment agree: (1) the plasma potential has been sustained at values near the discharge voltage, and (2) the electron temperature has been lowered by at least 2.5-3 times compared to the unshielded (US) thruster. Also, based on carbon deposition measurements, the erosion rates at the inner and outer walls of the MS thruster are found to be lower by at least 2300 and 1875 times, respectively. Erosion was so low along these walls that the rates were below the resolution of the profilometer. Using a sputtering yield model with an energy threshold of 25 V, the simulations predict a reduction of ~600 at the MS inner wall. At the outer wall ion energies are computed to be below 25 V, for which case we set the erosion to zero in the simulations. When a 50-V threshold is used the computed ion energies are below the threshold at both sides of the channel. Uncertainties, sensitivities and differences between theory and experiment are also discussed.

plasma simulations↗
NASA NTRS2008

Qualification of Commercial XIPS(R) Ion Thrusters for NASA Deep Space Missions

Electric propulsion systems based on commercial ion and Hall thrusters have the potential for significantly reducing the cost and schedule-risk of Ion Propulsion Systems (IPS) for deep space missions. The large fleet of geosynchronous communication satellites that use solar electric propulsion (SEP), which will approach 40 satellites by year-end, demonstrates the significant level of technical maturity and spaceflight heritage achieved by the commercial IPS systems. A program to delta-qualify XIPS(R) ion thrusters for deep space missions is underway at JPL. This program includes modeling of the thruster grid and cathode life, environmental testing of a 25-centimeter electromagnetic (EM) thruster over DAWN-like vibe and temperature profiles, and wear testing of the thruster cathodes to demonstrate the life and benchmark the model results. This paper will present the delta-qualification status of the XIPS thruster and discuss the life and reliability with respect to known failure mechanisms.

electric propulsion↗
NASA NTRS2015

Experimental Investigation of the Near-Wall Region in the NASA HiVHAc EDU2 Hall Thruster

The HiVHAc propulsion system is currently being developed to support Discovery-class NASA science missions. Presently, the thruster meets the required operational lifetime by utilizing a novel discharge channel replacement mechanism. As a risk reduction activity, an alternative approach is being investigated that modifies the existing magnetic circuit to shift the ion acceleration zone further downstream such that the magnetic components are not exposed to direct ion impingement during the thruster's lifetime while maintaining adequate thruster performance and stability. To measure the change in plasma properties between the original magnetic circuit configuration and the modified, "advanced" configuration, six Langmuir probes were flush-mounted within each channel wall near the thruster exit plane. Plasma potential and electron temperature were measured for both configurations across a wide range of discharge voltages and powers. Measurements indicate that the upstream edge of the acceleration zone shifted downstream by as much as 0.104 channel lengths, depending on operating condition. The upstream edge of the acceleration zone also appears to be more insensitive to operating condition in the advanced configuration, remaining between 0.136 and 0.178 channel lengths upstream of the thruster exit plane. Facility effects studies performed on the original configuration indicate that the plasma and acceleration zone recede further upstream into the channel with increasing facility pressure. These results will be used to inform further modifications to the magnetic circuit that will provide maximum protection of the magnetic components without significant changes to thruster performance and stability.

Electric Propulsion↗
NASA NTRS2015

Near-Surface Plasma Characterization of the 12.5-kW NASA TDU1 Hall Thruster

To advance the state-of-the-art in Hall thruster technology, NASA is developing a 12.5-kW, high-specific-impulse, high-throughput thruster for the Solar Electric Propulsion Technology Demonstration Mission. In order to meet the demanding lifetime requirements of potential missions such as the Asteroid Redirect Robotic Mission, magnetic shielding was incorporated into the thruster design. Two units of the resulting thruster, called the Hall Effect Rocket with Magnetic Shielding (HERMeS), were fabricated and are presently being characterized. The first of these units, designated the Technology Development Unit 1 (TDU1), has undergone extensive performance and thermal characterization at NASA Glenn Research Center. A preliminary lifetime assessment was conducted by characterizing the degree of magnetic shielding within the thruster. This characterization was accomplished by placing eight flush-mounted Langmuir probes within each discharge channel wall and measuring the local plasma potential and electron temperature at various axial locations. Measured properties indicate a high degree of magnetic shielding across the throttle table, with plasma potential variations along each channel wall being less than or equal to 5 eV and electron temperatures being maintained at less than or equal to 5 eV, even at 800 V discharge voltage near the thruster exit plane. These properties indicate that ion impact energies within the HERMeS will not exceed 26 eV, which is below the expected sputtering threshold energy for boron nitride. Parametric studies that varied the facility backpressure and magnetic field strength at 300 V, 9.4 kW, illustrate that the plasma potential and electron temperature are insensitive to these parameters, with shielding being maintained at facility pressures 3X higher and magnetic field strengths 2.5X higher than nominal conditions. Overall, the preliminary lifetime assessment indicates a high degree of shielding within the HERMeS TDU1, effectively mitigating discharge channel erosion as a life-limiting mechanism.

Lifetime↗
NASA NTRS2015

Magnesium Hall Thruster

This Phase II project is developing a magnesium (Mg) Hall effect thruster system that would open the door for in situ resource utilization (ISRU)-based solar system exploration. Magnesium is light and easy to ionize. For a Mars- Earth transfer, the propellant mass savings with respect to a xenon Hall effect thruster (HET) system are enormous. Magnesium also can be combusted in a rocket with carbon dioxide (CO2) or water (H2O), enabling a multimode propulsion system with propellant sharing and ISRU. In the near term, CO2 and H2O would be collected in situ on Mars or the moon. In the far term, Mg itself would be collected from Martian and lunar regolith. In Phase I, an integrated, medium-power (1- to 3-kW) Mg HET system was developed and tested. Controlled, steady operation at constant voltage and power was demonstrated. Preliminary measurements indicate a specific impulse (Isp) greater than 4,000 s was achieved at a discharge potential of 400 V. The feasibility of delivering fluidized Mg powder to a medium- or high-power thruster also was demonstrated. Phase II of the project evaluated the performance of an integrated, highpower Mg Hall thruster system in a relevant space environment. Researchers improved the medium power thruster system and characterized it in detail. Researchers also designed and built a high-power (8- to 20-kW) Mg HET. A fluidized powder feed system supporting the high-power thruster was built and delivered to Busek Company, Inc.

Szabo, James J.↗
NASA NTRS2015

Direct Drive Solar-Powered Arcjet Thruster

Electric thrusters typically require a power processing unit (PPU) to convert the spacecraft-provided power to the voltage and current that a thruster needs for operation. NASA Marshall Space Flight Center has initiated fundamental studies on whether an arcjet thruster can be operated directly with the power produced by solar arrays without any additional conversion. Elimination of the PPU significantly reduces system-level complexity of the propulsion system, and lowers developmental cost and risk. The proposed work will aim to refine the proof-of-concept presently being assembled and begin to identify and address technical questions related to power conditioning and noise suppression in the system, and heating of the thruster in long-duration operation. The apparatus proposed for investigation has a target power level of 400 to 1,000 W. The proposed direct-drive arcjet is potentially a highly scalable concept, applicable to spacecraft with up to hundreds of kilowatts and beyond. The design of the arcjet built for this effort was based on previous low power (1 kW class) arcjets.1-3 It has a precision machined 99.95% pure tungsten anode that also serves as the nozzle with a 0.040-in- (1-mm-) diameter, 0.040-in-long constrictor region. An additional anode with a 0.020-in- (0.5-mm-) diameter, 0.020-inlong constrictor region was purchased, but has not yet been used. The cathode is a 0.125-in-diameter tungsten welding electrode doped with lanthum-oxygen; its tip was precision ground to a 308deg angle and terminates in a blunt end. The two electrodes are separated by a boron-nitride insulator that also serves as the propellant manifold; it ends in six small holes which introduce the propellant gas in the diverging section of the nozzle, directly adjacent to the cathode. The electrodes and insulator are housed in a stainless-steel outer body, with a Macor insulator at the mid-plane to provide thermal isolation between the front and back halves of the device. The gas seals were made using Grafoil gaskets. Figure 1(a) shows the assembled thruster; figure 1(b) shows the thruster in the vacuum chamber with electrical and propellant connections.

Polzin, Kurt↗
NASA NTRS2016

Micropulsed Plasma Thrusters for Attitude Control of a Low-Earth-Orbiting CubeSat

This study presents a 3-Unit CubeSat design with commercial-off-the-shelf hardware, Teflon-fueled micropulsed plasma thrusters, and an attitude determination and control approach. The micropulsed plasma thruster is sized by the impulse bit and pulse frequency required for continuous compensation of expected maximum disturbance torques at altitudes between 400 and 1000 km, as well as to perform stabilization of up to 20 deg /s and slew maneuvers of up to 180 deg. The study involves realistic power constraints anticipated on the 3-Unit CubeSat. Attitude estimation is implemented using the q method for static attitude determination of the quaternion using pairs of the spacecraft-sun and magnetic-field vectors. The quaternion estimate and the gyroscope measurements are used with an extended Kalman filter to obtain the attitude estimates. Proportional-derivative control algorithms use the static attitude estimates in order to calculate the torque required to compensate for the disturbance torques and to achieve specified stabilization and slewing maneuvers or combinations. The controller includes a thruster-allocation method, which determines the optimal utilization of the available thrusters and introduces redundancy in case of failure. Simulation results are presented for a 3-Unit CubeSat under detumbling, pointing, and pointing and spinning scenarios, as well as comparisons between the thruster-allocation and the paired-firing methods under thruster failure.

Gatsonis, Nikolaos A.↗
NASA NTRS2018

Green Propellant Infusion Mission (GPIM) Thruster Valve Testing

The Green Propellant Infusion Mission (GPIM) is a project in the NASA Technology Demonstration Missions (TDM) that will perform a spaceflight demonstration of an AF-M315E propulsion system. The mission is led by Ball Aerospace & Technologies Corporation, with participation across industry and government. The spacecraft is based on the Ball BCP-100 bus, with an Aerojet propulsion system. NASA GSFC provides system testing and analysis of the flow within the AF-M315E propulsion system. A facility was developed at the NASA GSFC propulsion test site to safely use and test the AF-M315E propellant. Testing was performed on multiple 1 N and 22 N non-flight thruster valves provided by Aerojet. A mock-up of the flight system was driven at multiple pressures and temperatures to approximate different flight conditions. Testing was conducted to observe how AF-M315E behaves in fight-like conditions due to the propellant's mass properties heavy dependence on changes in temperature. The testing consisted of driving the different thruster valves at a set on-time and off-time for all and varying the driving pressure from 125 psig to 410 psig. Individual pulses were measured and analyzed for each valve to compare the various rise times and fall times of each thruster valve. The rise times and fall times were then analyzed for thruster valves by measuring the differential pressure from directly upstream and downstream of the valve. The rise times and fall times were also analyzed by measuring the amperage trace of the thruster valve measured with an oscilloscope. The pressure and amperage rise and fall times were compared collectively and individually for the various thruster valves. These results were compared to the fall time data provided by Aerojet.

Hynous, Andrew T.↗
NASA NTRS2021

Application of Micro-thrusters for Space Observatory Precision Attitude Control

This paper describes the results of a NASA investigation into the benefits of micro-thrusters compared to reaction wheels on future observatory-class missions with tight pointing stability requirements. Pointing repeatability and stability (i.e., jitter) requirements are key for space telescope missions of the future. For example, managing jitter is essential to being able to “image” planets orbiting distant stars. Jitter requirements for missions in this class are difficult to meet with current reaction wheel-based architectures. The reaction wheels are typically the largest pointing disturbance on the spacecraft. Disturbances from reaction wheels can be mitigated, typically by mechanically isolating the wheels, which imposes system complexity and cost. Thrusters capable of thrust forces in the micronewton (μN) range (referred to as micro-thrusters or micronewton thrusters) have been developed to support the Laser Interferometer Space Antenna (LISA) mission, which requires drag-free control to place a test mass in near-perfect free-fall Beyond the drag-free control application, micro-thrusters could be used as a substitute for reaction wheels or as a supplement to wheels for fine pointing control. Used in this fashion, micro-thrusters have potential for reducing the cost and technical risks of achieving demanding pointing stability performance on observatory-class missions.

Chiang, Richard Y.↗
NASA NTRSDate not supplied

Autonomous Control for Arbitrary Thruster Configurations and Mass Properties in Special Euclidean Group SE(3)

Most current methods for determining maneuvers and thrust firing sequences depend on explicit and predetermined commands generated by a combination of on-board systems and ground-based human-in-the-loop methods. For spacecraft and space structures with changing mass properties and thruster configurations, such as the Deep Space Gateway as it changes configurations throughout its lifetime, determining these commands can be time-consuming and computationally intensive. However, recent work within the Lie GroupSE (3) has offered ways of autonomously determining the location, power, precision, and capabilities of thrusters in any arbitrary position. Furthermore, a method for determining thruster firing sequences based on an arbitrary control input (both translational and rotational in a coupled, 6-element vector) and arbitrary thruster configurations has also recently been developed. When combining these methods, any spacecraft with any mass properties and thruster configurations can be understood in terms of controllability limits and thruster firing sequences can be generated quickly and with low computational load, thus extending the autonomous capabilities of deep space missions. In this work, this method is presented and explored in terms of computational load, robustness in the presence of uncertainty, and overall performance. The capabilities of this method are also examined in the case of the Deep Space Gateway both in fully controllable configurations and uncontrollable configurations.

Control↗
NASA NTRSDate not supplied

Autonomous Control for Arbitrary Thruster Configurations and Mass Properties in Special Euclidean Group SE(3)

Most current methods for determining maneuvers and thrust firing sequences depend on explicit and predetermined commands generated by a combination of on-board systems and ground-based human-in-the-loop methods. For spacecraft and space structures with changing mass properties and thruster configurations, such as the Deep Space Gateway as it changes configurations throughout its lifetime, determining these commands can be time-consuming and computationally intensive. However, recent work within the Lie group SE(3) has offered ways of autonomously determining the location, power, precision, and capabilities of thrusters in any arbitrary position. Furthermore, a method for determining thruster firing sequences based on an arbitrary control input (both translational and rotational in a coupled, 6-element vector) and arbitrary thruster configurations has also recently been developed. When combining these methods, any spacecraft with any mass properties and thruster configurations can be understood in terms of controllability limits and thruster firing sequences can be generated quickly and with low computational load, thus extending the autonomous capabilities of deep space missions. In this work, this method is presented and explored in terms of convergence time to the desired pose. The capabilities of this method are also examined in the case of the Deep Space Gateway both in fully controllable configurations and uncontrollable configurations.

SE(3)↗
DOE OSTI2013

Radial scale effect on the performance of low-power cylindrical Hall plasma thrusters

Investigation of the radial scale effect on low-power cylindrical Hall thrusters has been undertaken by comparing the thrusters with three different channel diameters of 28, 40, and 50 mm. The investigation found that both the anode efficiency and the thrust of the larger thruster are higher as the anode power is raised. On the other hand, higher current and propellant utilizations are achieved for the smaller thruster, which is due to higher neutral density and better electron confinement. The large plume angle of the small cylindrical Hall thruster causes thrust loss, resulting in the reduction of anode efficiency.

70 PLASMA PHYSICS AND FUSION TECHNOLOGY↗
DOE OSTI2014

Magnetic shielding of a laboratory Hall thruster. I. Theory and validation

We demonstrate a technique by which erosion of the acceleration channel in Hall thrusters can be reduced by at least a few orders of magnitude. The first principles of the technique, now known as “magnetic shielding,” have been derived based on the findings of 2-D numerical simulations. The simulations, in turn, guided the modification of an existing 6-kW laboratory Hall thruster to test the theory and are the main subject of this Part I article. Part II expands on the results of the experiments. Near the walls of the magnetically shielded (MS) thruster theory and experiment agree that (1) the plasma potential has been sustained at values near the discharge voltage, and (2) the electron temperature has been lowered compared to the unshielded thruster. Erosion rates deduced directly from the wall probes show reductions of at least ∼3 orders of magnitude at the MS inner wall when an ion energy threshold of 30.5 V is used in the sputtering yield model of the channel material. At the outer wall the probes reveal that the ion energy was below the assumed threshold. Using a threshold of 25 V, the simulations predict a minimum reduction of ∼600 at the MS inner wall. At the MS outer wall ion energies are found to be below 25 V. When a 50-V threshold is used the computed ion energies are below the threshold at both sides of the channel. Uncertainties, sensitivities, and differences between theory and experiment are also discussed. The elimination of wall erosion in Hall thrusters solves a problem that has remained unsettled for more than five decades.

70 PLASMA PHYSICS AND FUSION TECHNOLOGY↗
DOE OSTI2014

Sheath oscillation characteristics and effect on near-wall conduction in a krypton Hall thruster

Despite its affordability, the krypton Hall-effect thruster in applications always had problems in regard to performance. The reason for this degradation is studied from the perspective of the near-wall conductivity of electrons. Using the particle-in-cell method, the sheath oscillation characteristics and its effect on near-wall conduction are compared in the krypton and xenon Hall-effect thrusters both with wall material composed of BNSiO{sub 2}. Comparing these two thrusters, the sheath in the krypton-plasma thruster will oscillate at low electron temperatures. The near-wall conduction current is only produced by collisions between electrons and wall, thereby causing a deficiency in the channel current. The sheath displays spatial oscillations only at high electron temperature; electrons are then reflected to produce the non-oscillation conduction current needed for the krypton-plasma thruster. However, it is accompanied with intensified oscillations.

70 PLASMA PHYSICS AND FUSION TECHNOLOGY↗
DOE OSTI2016

Operating a magnetic nozzle helicon thruster with strong magnetic field

A pulsed axial magnetic field up to ∼2.8 kG is applied to a 26-mm-inner-diameter helicon plasma thruster immersed in a vacuum chamber, and the thrust is measured using a pendulum target. The pendulum is located 30-cm-downstream of the thruster, and the thruster rf power and argon flow rate are fixed at 1 kW and 70 sccm (which gives a chamber pressure of 0.7 mTorr). The imparted thrust increases as the applied magnetic field is increased and saturates at a maximum value of ∼9.5 mN for magnetic field above ∼2 kG. At the maximum magnetic field, it is demonstrated that the normalized plasma density, and the ion flow energy in the magnetic nozzle, agree within ∼50% and of 10%, respectively, with a one-dimensional model that ignores radial losses from the nozzle. This magnetic nozzle model is combined with a simple global model of the thruster source that incorporates an artificially controlled factor α, to account for radial plasma losses to the walls, where α = 0 and 1 correspond to zero losses and no magnetic field, respectively. Comparison between the experiments and the model implies that the radial losses in the thruster source are experimentally reduced by the applied magnetic field to about 10% of that obtained from the no magnetic field model.

70 PLASMA PHYSICS AND FUSION TECHNOLOGY↗