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NASA NTRS2014

Mode Transitions in Magnetically Shielded Hall Effect Thrusters

A mode transition study is conducted in magnetically shielded thrusters where the magnetic field magnitude is varied to induce mode transitions. Three different oscillatory modes are identified with the 20-kW NASA-300MS-2 and the 6-kW H6MS: Mode 1) global mode similar to unshielded thrusters at low magnetic fields, Mode 2) cathode oscillations at nominal magnetic fields, and Mode 3) combined spoke, cathode and breathing mode oscillations at high magnetic fields. Mode 1 exhibits large amplitude, low frequency (1-10 kHz), breathing mode type oscillations where discharge current mean value and oscillation amplitude peak. The mean discharge current is minimized while thrust-to-power and anode efficiency are maximized in Mode 2, where higher frequency (50-90 kHz), low amplitude, cathode oscillations dominate. Thrust is maximized in Mode 3 and decreases by 5-6% with decreasing magnetic field strength. The presence or absence of spokes and strong cathode oscillations do not affect each other or discharge current. Similar to unshielded thrusters, mode transitions and plasma oscillations affect magnetically shielded thruster performance and should be characterized during system development.

Ground Tests↗

NASA NTRS2009

Discharge Oscillations in a Permanent Magnet Cylindrical Hall-Effect Thruster

Measurements of the discharge current in a cylindrical Hall thruster are presented to quantify plasma oscillations and instabilities without introducing an intrusive probe into the plasma. The time-varying component of the discharge current is measured using a current monitor that possesses a wide frequency bandwidth and the signal is Fourier transformed to yield the frequency spectra present, allowing for the identification of plasma oscillations. The data show that the discharge current oscillations become generally greater in amplitude and complexity as the voltage is increased, and are reduced in severity with increasing flow rate. The breathing mode ionization instability is identified, with frequency as a function of discharge voltage not increasing with discharge voltage as has been observed in some traditional Hall thruster geometries, but instead following a scaling similar to a large-amplitude, nonlinear oscillation mode recently predicted in for annular Hall thrusters. A transition from lower amplitude oscillations to large relative fluctuations in the oscillating discharge current is observed at low flow rates and is suppressed as the mass flow rate is increased. A second set of peaks in the frequency spectra are observed at the highest propellant flow rate tested. Possible mechanisms that might give rise to these peaks include ionization instabilities and interactions between various oscillatory modes.

Polzin, K. A.↗

NASA NTRS2017

High Current Lanthanum Hexaboride Hollow Cathode for 20-200 kW Hall Thrusters

The demand for higher power Hall and ion thrusters continues for providing high thrust and long life for deep space mission. The X3 nested Hall thruster is capable of power levels in the 20 to 200-kW range, and is being developed by the University of Michigan and Aerojet for future cargo and manned-missions. The cathode for this thruster is required to produce discharge currents of 50 to 350 A with lifetimes in excess of 10 khrs. A high-current lanthanum hexaboride (LaB6) hollow cathode was previously developed at JPL for these applications, and was successfully operated at over 250 A of discharge current. An updated version of this cathode has been designed, built and tested at JPL, and then used to run the X3 nested Hall thruster at currents of up to 250 A. The new version is designed to reduced orifice plate overheating at high currents, and is capable of injecting auxiliary gas directly into the near-cathode plume from two locations to minimize energetic ion generation at high current. The cathode is predicted to be capable of producing over 350 A of discharge current, and has been tested to date at JPL at steady-state discharge currents from 25 A to 300 A.

Goebel, Dan M.↗

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 NTRS1978

Effect of anode-cathode geometry on performance of the HIP-1 hot ion plasma

Hot-ion hydrogen plasma experiments were conducted in the NASA Lewis HIP-1 magnetic mirror facility to determine how the ion temperature was influenced by the axial position of the cathode tips relative to the anodes. A steady-state EXB plasma was formed by applying a strong radially inward dc electric field near the throats of the magnetic mirrors. The dc electric field was created between hollow cathode rods inside hollow anode cylinders, both concentric with the magnetic axis. The highest ion temperatures, 900 eV, were attained when the tip of each cathode was in the same plane as the end of its anode. These temperatures were reached with 22 kV applied to the electrodes in a field of 1.1 tesla. Scaling relations were empirically determined for ion temperature and the product of ion density and neutral particle density as a function of cathode voltage, discharge current, and electrode positions. Plasma discharge current vs voltage (I-V) characteristics were determined.

Lauver, M. R.↗

NASA NTRS1991

Preliminary test results of a hollow cathode MPD thruster

Performance of four hollow cathode configurations with low work function inserts was evaluated in a steady-state 100 kW class applied magnetic field magnetoplasmadynamic (MPD) thruster. Two of the configurations exhibited stable discharge current attachment to the low work function inserts of the hollow cathodes. A maximum discharge current of 2250 A was attained. While the applied-field increased the performance of the thruster, at high applied fields the discharge current attachment moved from the insert to the cathode body. The first successful hollow cathode performed well in comparison with a conventional rod cathode MPD thruster, attaining a thrust efficiency with argon of close to 20 percent at a specific impulse of about 2000 s. The second successful configuration had significantly lower performance.

Mantenieks, Maris A.↗

NASA NTRS2009

Gallium Electromagnetic (GEM) Thruster Performance Measurements

Discharge current, terminal voltage, and mass bit measurements are performed on a coaxial gallium electromagnetic thruster at discharge currents in the range of 7-23 kA. It is found that the mass bit varies quadratically with the discharge current which yields a constant exhaust velocity of 20 km/s. Increasing the electrode radius ratio of the thruster from to 2.6 to 3.4 increases the thruster efficiency from 21% to 30%. When operating with a central gallium anode, macroparticles are ejected at all energy levels tested. A central gallium cathode ejects macroparticles when the current density exceeds 3.7 10(exp 8) A/square m . A spatially and temporally broad spectroscopic survey in the 220-520 nm range is used to determine which species are present in the plasma. The spectra show that neutral, singly, and doubly ionized gallium species are present in the discharge, as well as annular electrode species at higher energy levels. Axial Langmuir triple probe measurements yield electron temperatures in the range of 0.8-3.8 eV and electron densities in the range of 8 x 10(exp )20 to 1.6 x 10(exp 21) m(exp -3) . Triple probe measurements suggest an exhaust plume with a divergence angle of 9 , and a completely doubly ionized plasma at the ablating thruster cathode.

Thomas, Robert E.↗

NASA NTRS2010

Investigation of a Gallium MPD Thruster with an Ablating Cathode

Arc impedance, exhaust velocity, and plasma probe measurements are presented. The thruster is driven by a 50 microsecond pulse from a 6.2 milliohm pulse forming network, and gallium is supplied to the discharge by evaporation of the cathode. The arc voltage is found to vary linearly with the discharge current with an arc impedance of 6.5 milliohms. Electrostatic probes yield an exhaust velocity that is invariant with the discharge current and has a peak value of 20 kilometers per second, which is in reasonable agreement with the value (16 plus or minus 1 kilometer per second) calculated from the mass bit and discharge current data. Triple probe measurements yield on axis electron temperatures in the range of 0.8-3.8 eV, electron densities in the range of 1.6 x 10(exp 21) to 2.1 x 10(exp 22) per cubic meter, and a divergence half angle of 16 degrees. Measurements within the interelectrode region yield a peak magnetic field of 0.8 T, and the observed radial trends are consistent with an azimuthally symmetric current distribution. A cathode power balance model is coupled with an ablative heat conduction model predicting mass bit values that are within 20% of the experimental values.

Thomas, Robert E.↗

NASA NTRS2014

Investigation of Energetic Ions in a 100-A Hollow Cathode

The role of ion acoustic turbulence in the formation of high-energy ion tails in the plume of a 100-A LaB6 hollow cathode is experimentally and theoretically examined. At fixed flow rate and varying discharge current, single-point measurements of fluctuation intensity in the cathode plume are taken and compared to ion energy measurements. It is shown that for high discharge current the formation of energetic ions is correlated with the amplitude of the ion acoustic turbulence. Two-dimensional maps of background plasma parameters and wave turbulence are made at the highest discharge current investigated, 140 A. A simple, one-dimensional quasilinear model for the interaction of the ion energy distribution with the ion acoustic turbulence is employed, and it is shown that the energy in the measured wave turbulence is sufficiently large to explain the formation of ion tails in the cathode plume. Mitigation techniques for minimizing the amplitude of the turbulence are discussed.

ion acoustic turbulence↗

NASA NTRS1985

Characteristics of EMI generated by negative metal-positive dielectric voltage stresses due to spacecraft charging

Charging of spacecraft surfaces by the environmental plasma can result in differential potentials between metallic structure and adjacent dielectric surfaces in which the relative polarity of the voltage stress is either negative dielectric/positive metal or negative metal/positive dielectric. Negative metal/positive dielectric is a stress condition that may arise if relatively large areas of spacecraft surface metals are shadowed from solar UV and/or if the UV intensity is reduced as in the situation in which the spacecraft is entering into or leaving eclipse. The results of experimental studies of negative metal/positive dielectric systems are given. Information is given on: enhanced electron emission I-V curves; e(3) corona noise vs e(3) steady-state current; the localized nature of e(3) and negative metal arc discharge currents; negative metal arc discharges at stress thresholds below 1 kilovolt; negative metal arc discharge characteristics; dependence of blowoff arc discharge current on spacecraft capacitance to space (linear dimension); and damage to second surface mirrors due to negative metal arcs.

Chaky, R. C.↗

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 NTRS1977

A hollow cathode hydrogen ion source

High current density ion sources have been used to heat plasmas in controlled thermonuclear reaction experiments. High beam currents imply relatively high emission currents from cathodes which have generally taken the form of tungsten filaments. This paper describes a hydrogen ion source which was primarily developed to assess the emission current capability and design requirements for hollow cathodes for application in neutral injection devices. The hydrogen source produced ions by electron bombardment via a single hollow cathode. Source design followed mercury ion thruster technology, using a weak magnetic field to enhance ionization efficiency. A 1.3-cm-diam hollow cathode using a low work function material dispenser performed satisfactorily over a discharge current range of 10-90 A. Cylindrical probe measurements taken without ion extraction indicate maximum ion number densities on the order of 10 trillion/cu cm. Discharge durations ranged from 30 sec to continuous operation. Tests with beam extraction at 2.5 keV and 30 A discharge current yield average ion beam current densities of 0.1 A/sq cm over a 5-cm extraction diameter. Results of this study can be used to supply the baseline information needed to scale hollow cathodes for operation at discharge currents of hundreds of amperes using distributed cathodes.

Sovey, J. S.↗

NASA NTRS1992

Scaling of 100 kW class applied-field MPD thrusters

Three cylindrical applied-field magnetoplasmadynamic thrusters were tested with argon propellant over a broad range of operating conditions to establish empirical scaling laws for thruster performance. Argon flow rates, discharge currents, and applied-field strengths were varied between 0.025 and 0.14 g/s, 750 to 2000 A, and 0.034 to 0.20 T, respectively. The results showed that the thrust reached over five times the self-field value, and that thrust increased linearly with the product of discharge current and applied-field strength, and quadratically with the anode radius. While increasing the propellant flow rate increased the thrust, it did not affect the rate of thrust increase with applied-field strength, and, at low propellant flow rates, the self-field thrust approached 30 percent of the measured thrust. The voltage increased linearly with applied-field strength but was insensitive to the discharge current. The rate of voltage increase with applied-field strength was strongly dependent on anode radius. Thruster efficiency increased monotonically with applied-field strength and propellant flow rate.

Myers, Roger M.↗

NASA NTRS1992

Scaling of 100 kW class applied-field MPD thrusters

Three cylindrical applied-field magnetoplasmadynamic thrusters were tested with argon propellant over a broad range of operating conditions to establish empirical scaling laws for thruster performance. Argon flow rates, discharge currents, and applied-field strengths were varied between 0.025 and 0.14 g/s, 750 to 2000 A, and 0.034 to 0.20T, respectively. The results showed that the thrust reached over five times the self-field value, and that thrust increased linearly with the product of discharge current and applied-field strength and quadratically with the anode radius. While increasing the propellant flow rate increased the thrust, it did not affect the rate of thrust increase with applied-field strength, and at low propellant flow rates the self-field thrust approached 30 percent of the measured thrust. The voltage increased linearly with applied-field strength but was insensitive to the discharge current. The rate of voltage increase with applied-field strength dependent on anode radius. Thruster efficiency increased monotonically with applied-field strength and propellant flow rate. Peak thruster efficiencies were insensitive to changes in anode radius. Electrode power loss and thruster efficiency measurements showed that while the electrode losses decreased with increasing anode radius the plasma losses increased. The opposite behaviors of electrode and plasma losses demonstrates the need to identify ways of independently controlling the thruster loss mechanisms.

Myers, Roger M.↗

NASA NTRS1989

Plume characteristics of MPD thrusters - A preliminary examination

A diagnostics facility for MPD thruster plume measurements has been built and is currently undergoing testing. The facility includes electrostatic probes for electron temperature and density measurements, Hall probes for magnetic field and current distribution mapping, and an imaging system to establish the global distribution of plasma species. Preliminary results for MPD thrusters operated at power levels between 30 and 60 kW with solenoidal applied magnetic fields show that the electron density decreases exponentially from 1 x 10 to the 20th/cu m to 2 x 10 to the 18th/cu m over the first 30 cm of the expansion, while the electron temperature distribution is relatively uniform, decreasing from approximately 2.5 eV to 1.5 eV over the same distance. The radiant intensity of the ArII 4879 A line emission follows similar patterns. Current distribution measurements indicate that significant fraction of the discharge current is blown into the plume region, and that its distribution depends on the magnitudes of both the discharge current and the applied magnetic field.

Myers, R. M.↗

NASA NTRS2013

Optical Emission Characterization of High-Power Hall Thruster Wear

Optical emission spectroscopy is employed to correlate BN insulator erosion with high-power operation of the NASA 300M Hall-effect thruster. 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, discharge current and magnetic field strength. The boron signals are shown to trend with discharge current and show weak dependence on discharge voltage. The trends are consistent with data previously collected on the NASA 300M and NASA 457M thrusters but are different from conventional wisdom.

WIlliams, George J.↗

NASA NTRS1989

Plume characteristics of MPD thrusters: A preliminary examination

A diagnostics facility for MPD thruster plume measurements was built and is currently undergoing testing. The facility includes electrostatic probes for electron temperature and density measurements, Hall probes for magnetic field and current distribution mapping, and an imaging system to establish the global distribution of plasma species. Preliminary results for MPD thrusters operated at power levels between 30 and 60 kW with solenoidal applied magnetic fields show that the electron density decreases exponentially from 1x10(2) to 2x10(18)/cu m over the first 30 cm of the expansion, while the electron temperature distribution is relatively uniform, decreasing from approximately 2.5 eV to 1.5 eV over the same distance. The radiant intensity of the ArII 4879 A line emission also decays exponentially. Current distribution measurements indicate that a significant fraction of the discharge current is blown into the plume region, and that its distribution depends on the magnitudes of both the discharge current and the applied magnetic field.

Myers, Roger M.↗

NASA NTRS1983

Discharge characteristics of a simulated solar cell array

Measurements of the transient characteristics associated with the discharge of a simulated solar array are presented. A capacitively coupled probe is used to measure the discharge current, and antennas are used to measure the electromagnetic radiation. Discharges were observed at low surface voltages and several modes of discharge were observed. The maximum discharge current is found to be 0.2 A. This value is several orders of magnitude higher than that reported by previous measurements. Experimental evidence suggests that the inverted voltage gradient is a very likely triggering mechanism for solar array discharges.

Leung, P.↗