SEARCH · Search NASA

Results for “cathode plume”

Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

SourceResearch connectionFrom yearThrough year

Quote a phrase for an exact phrase match. Source license links do not imply unrestricted reuse.

At least 55 records · Page 3

All sources NASA NTRS DOE OSTI

NASA NTRS1978

An investigation of mercury hollow cathode phenomena

An experimental investigation is described in which measurements were made of plasma properties and other physical parameters upstream and downstream of the orifice plate of an operating mercury hollow cathode. These measurements included plasma property profiles, internal surface temperatures, internal pressure, and spectroscopic data of the external discharge. The results show the plume-to-spot mode transition to be characterized by significant changes in the plasma properties downstream of the cathode orifice, including the disappearance of an apparent electron free-fall region which is found to be present in the plume mode. The results also indicate the existence of a very dense plasma (10 to the 10th to 10 to the 14th electrons/cu cm) with steep electron density gradients in the region upstream of the orifice. The upstream data is found to correlate well with a modified form of the Saha equation. The related analysis indicates that the heavy particles approach thermal equilibrium with the electrons in the region of the orifice and that the ionization fraction in the orifice is very high - 80 to 100%.

Siegfried, D. E.↗

NASA NTRS2006

Near Discharge Cathode Assembly Plasma Potential Measurements in a 30-cm NSTAR Type Ion Engine During Beam Extraction

Floating emissive probe plasma potential data are presented over a two-dimensional array of locations in the near Discharge Cathode Assembly (DCA) region of a 30-cm diameter ring-cusp ion thruster. Discharge plasma data are presented with beam extraction at throttling conditions comparable to the NASA TH Levels 8, 12, and 15. The operating conditions of the Extended Life Test (ELT) of the Deep Space One (DS1) flight spare ion engine, where anomalous discharge keeper erosion occurred, were TH 8 and TH 12 consequently they are of specific interest in investigating discharge keeper erosion phenomena. The data do not validate the presence of a potential hill plasma structure downstream of the DCA, which has been proposed as a possible erosion mechanism. The data are comparable in magnitude to data taken by other researchers in ring-cusp electron-bombardment ion thrusters. The plasma potential structures are insensitive to thruster throttling level with a minimum as low as 14 V measured at the DCA exit plane and increasing gradually in the axial direction. A sharp increase in plasma potential to the bulk discharge value of 26 to 28 volts, roughly 10 mm radially from DCA centerline, was observed. Plasma potential measurements indicate a low-potential plume structure that is roughly 20 mm in diameter emanating from the discharge cathode that may be attributed to a free-standing plasma double layer.

Herman, Daniel A.↗

NASA NTRS1994

The NASA-JPL advanced propulsion program

The NASA Advanced Propulsion Concepts (APC) program at the Jet Propulsion Laboratory (JPL) consists of two main areas: The first involves cooperative modeling and research activities between JPL and various universities and industry; the second involves research at universities and industry that is directly supported by JPL. The cooperative research program consists of mission studies, research and development of ion engine technology using C-60 (Buckminsterfullerene) propellant, and research and development of lithium-propellant Lorentz-force accelerator (LFA) engine technology. The university/industry- supported research includes research (modeling and proof-of-concept experiments) in advanced, long-life electric propulsion, and in fusion propulsion. These propulsion concepts were selected primarily to cover a range of applications from near-term to far-term missions. For example, the long-lived pulsed-xenon thruster research that JPL is supporting at Princeton University addresses the near-term need for efficient, long-life attitude control and station-keeping propulsion for Earth-orbiting spacecraft. The C-60-propellant ion engine has the potential for good efficiency in a relatively low specific impulse (Isp) range (10,000 - 30,000 m/s) that is optimum for relatively fast (less than 100 day) cis-lunar (LEO/GEO/Lunar) missions employing near-term, high-specific mass electric propulsion vehicles. Research and modeling on the C-60-ion engine are currently being performed by JPL (engine demonstration), Caltech (C-60 properties), MIT (plume modeling), and USC (diagnostics). The Li-propellant LFA engine also has good efficiency in the modest Isp range (40,000 - 50,000 m/s) that is optimum for near-to-mid-term megawatt-class solar- and nuclear-electric propulsion vehicles used for Mars missions transporting cargo (in support of a piloted mission). Research and modeling on the Li-LFA engine are currently being performed by JPL (cathode development), Moscow Aviation Institute (engine testing), Thermacore (electrode development), as well as at MIT (plume modeling), and USC (diagnostics). Also, the mission performance of a nuclear-electric propulsion (NEP) Li-LFA Mars cargo vehicle is being modeled by JPL (mission analysis; thruster and power processor modeling) and the Rocketdyne Energy Technology and Engineering Center (ETEC) (power system modeling). Finally, the fusion propulsion research activities that JPL is supporting at Pennsylvania State University (PSU) and at Lawrenceville Plasma Physics (LPP) are aimed at far-term fast (less than 100 day round trip) piloted Mars missions and, in the very far term, interstellar missions.

Frisbee, Robert H.↗

NASA NTRS1992

Characterization of electric thruster plumes using multiplexed laser induced fluorescence measurements

The use of laser-induced fluorescence to obtain spatially resolved measurements of propellant velocities and temperatures in electric thruster plumes is discussed, with emphasis on two innovations of the technique, namely simultaneous recording of the optogalvanic signal in a hollow cathode lamp for the purpose of calibrating Doppler shifts, and two-beam multiplexing to allow the measurement of two velocity components at once. It is also shown how information on plume fluctuations can be obtained from the multiplxed LIF data. The techniques are demonstrated on the plume from a low power arcjet, operated on argon, and its extension to the measurement of ion velocities in electrostatic ion thrusters and stationary plasma thrusters is discussed.

Ruyten, W. M.↗

NASA NTRS1999

MEMS Micropropulsion Activities at JPL

A status of MEMS-based micropropulsion activities conducted at JPL will be given. These activities include work conducted on the so called Vaporizing Liquid Micro-Thruster (VLM) which recently underwent proof-of-concept testing, demonstrating the ability to vaporize water propellant at 2 W and 2 V. Micro-ion engine technologies, such m field emitter arrays and micro-grids are being studied. Focus in the field emitter area is on arrays able to survive in thruster plumes and micro-ion engine plasmas to serve as neutralizers aW engine cathodes. Integrated, batch-fabricated Ion repeller grid structures are being studied as well as different emitter tip materials are being investigated to meet these goals. A micro-isolation valve is being studied to isolate microspacecraft feed system during long interplanetary cruises, avoiding leakage and prolonging lifetime and reliability of such systems. This concept relies on the melting of a thin silicon barrier. Burst pressure values as high as 2,900 psig were obtained for these valves and power requirements to melt barriers ranging between 10 - 50 microns in thickness, as determined through thermal finite element calculations, varied between 10 - 30 W to be applied over a duration of merely 0.5 ms.

Mueller, Juergen↗

NASA NTRS1992

Absolute Doppler shift calibration of laser induced fluorescence signals using optogalvanic measurements in a hollow cathode lamp

The paper investigates the use of optogalvanic (OG) measurements on the neutral 3P1 and 3P2 levels of argon in a hollow cathode lamp for the purpose of calibrating Doppler shifts of laser-induced fluorescence signals from an arcjet plume. It is shown that, even with non-Doppler-free OG detection, accuracy to better than 10 MHz is possible but that, depending on the experiment geometry, corrections of 10-35 MHz may be necessary to offset small axial drift velocities of neutral atoms in the hollow cathode lamp.

Ruyten, Wilhelmus M.↗

NASA NTRS1987

Advanced electric propulsion and space plasma contactor research

A theory of the plasma contacting process is described and experimental results obtained using three different hollow cathode-based plasma contactors are presented. The existence of a sheath across which the bulk of the voltage drop associated with the contacting process occurs is demonstrated. Test results are shown to agree with a model of a spherical, space-charge-limited double sheath. The concept of ignited mode contactor operation is discussed, which is shown to enhance contactor efficiency when it is collecting electrons. An investigation of the potentials in the plasma plumes downstream of contactors operating at typical conditions is presented. Results of tests performed on hollow cathodes operating at high interelectrode pressures (up to about 1000 Torr) on ammonia are presented and criteria that are necessary to ensure that the cathode will operate properly in this regime are presented. These results suggest that high pressure hollow cathode operation is difficult to achieve and that special care must be taken to assure that the electron emission region remains diffuse and attached to the low work function insert. Experiments conducted to verify results obtained previously using a ring cusp ion source equipped with a moveable anode are described and test results are reported. A theoretical study of hollow cathode operation at high electron emission currents is presented. Preliminary experiments using the constrained sheath optics concept to achieve ion extraction under conditions of high beam current density, low net accelerating voltage and well columniated beamlet formation are discussed.

Wilbur, Paul J.↗

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 NTRS2019

The Effect of Anode Position on Operation of a 25-A Class Hollow Cathode

The effect of the distance between a hollow cathode and a cylindrical anode on cathode operation is investigated for two anode geometries. Neutral flow simulations demonstrate that the anode diameter and distance from the cathode exit can elevate the downstream pressure as much as two orders of magnitude above what the cathode experiences while operating within a Hall thruster. Based on the results of this modeling, two axially-segmented cylindrical molybdenum anodes were constructed: a 64-mm diameter one that replicated the anode geometry used in recent NASA hollow cathode development testing and a larger 254-mm diameter one designed to reduce the neutral pressure in front of the cathode to thruster-like values. For each anode design, cathode performance was characterized for varying anode/cathode distance using metrics such as discharge voltage and oscillation magnitudes, and the ion voltage spectra were characterized using a radially-positioned retarding potential analyzer. It was found that as local neutral pressure decreased, discharge voltage and high-voltage ion content in the plume increased. For the 254-mm diameter anode, an ion voltage tail in excess of 200 V was found for nominal cathode flow rates. The implications of these results for standalone hollow cathode development tests are discussed.

Hall, Scott J.↗

DOE OSTI2024

Establishing criteria for the transition from kinetic to fluid modeling in hollow cathode analysis

Hollow cathodes for plasma switch applications are investigated via 2D3V particle-in-cell simulations of the channel and plume region. The kinetic nature of the plasma within the channel is dependent on the thermalization rate of electrons, emitted from the insert. When Coulomb collisions occur at a much greater rate than ionization or excitation collisions, the electron energy distribution function rapidly relaxes to a Maxwellian and the plasma within the channel can be described accurately via a fluid model. In contrast, if inelastic processes are much faster than Coulomb collisions, then the electron energy distribution function in the channel exhibits a notable high-energy tail, and a kinetic treatment is required. This criterion is applied to hollow cathodes from the literature, revealing that a fluid approach is suitable for most electric propulsion applications, whereas a kinetic treatment can be more critical to accurate modeling of plasma switches.

70 PLASMA PHYSICS AND FUSION TECHNOLOGY↗

NASA NTRS1993

Plasma contactor development for Space Station

Plasma contactors have been baselined for the Space Station (SS) to control the electrical potentials of surfaces to eliminate/mitigate damaging interactions with the space environment. The system represents a dual-use technology which is a direct outgrowth of the NASA electric propulsion program and, in particular, the technology development effort on ion thrustor systems. The plasma contactor subsystems include the plasma contactor unit, a power electronics unit, and an expellant management unit. Under this pre-flight development program these will all be brought to breadboard or engineering model status. Development efforts for the plasma contactor include optimizing the design and configuration of the contactor, validating its required lifetime, and characterizing the contactor plume and electromagnetic interference. The plasma contactor unit design selected for the SS is an enclosed keeper, xenon hollow cathode plasma source. This paper discusses the test results and development status of the plasma contactor unit subsystem for the SS.

Patterson, Michael J.↗

NASA NTRS2009

Neutralizer Hollow Cathode Simulations and Comparisons with Ground Test Data

The fidelity of electric propulsion physics-based models depends largely on the validity of their predictions over a range of operating conditions and geometries. In general, increased complexity of the physics requires more extensive comparisons with laboratory data to identify the region(s) that lie outside the validity of the model assumptions and to quantify the uncertainties within its range of application. This paper presents numerical simulations of neutralizer hollow cathodes at various operating conditions and orifice sizes. The simulations were performed using a two-dimensional axisymmetric model that solves numerically a relatively extensive system of conservation laws for the partially ionized gas in these devices. A summary of the comparisons between simulation results and Langmuir probe measurements is provided. The model has also been employed to provide insight into recent ground test observations of the neutralizer cathode in NEXT. It is found that a likely cause of the observed keeper voltage drop is cathode orifice erosion. However, due to the small magnitude of this change, is approx. 0.5 V (less than 5% of the beginning-of-life value) over 10 khrs, and in light of the large uncertainties of the cathode material sputtering yield at low ion energies, other causes cannot be excluded. Preliminary simulations to understand transition to plume mode suggest that in the range of 3-5 sccm the existing 2-D model reproduces fairly well the rise of the keeper voltage in the NEXT neutralizer as observed in the laboratory. At lower flow rates the simulation produces oscillations in the keeper current and voltage that require prohibitively small time-steps to resolve with the existing algorithms.

electric propulsion↗

NASA NTRS1979

Inert gas thrusters

Inert gas thrusters considered for space propulsion systems were investigated. Electron diffusion across a magnetic field was examined utilizing a basic model. The production of doubly charged ions was correlated using only overall performance parameters. The use of this correlation is therefore possible in the design stage of large gas thrusters, where detailed plasma properties are not available. Argon hollow cathode performance was investigated over a range of emission currents, with the positions of the inert, keeper, and anode varied. A general trend observed was that the maximum ratio of emission to flow rate increased at higher propellant flow rates. It was also found that an enclosed keeper enhances maximum cathode emission at high flow rates. The maximum cathode emission at a given flow rate was associated with a noisy high voltage mode. Although this mode has some similarities to the plume mode found at low flows and emissions, it is encountered by being initially in the spot mode and increasing emission. A detailed analysis of large, inert-gas thruster performance was carried out. For maximum thruster efficiency, the optimum beam diameter increases from less than a meter at under 2000 sec specific impulse to several meters at 10,000 sec. The corresponding range in input power ranges from several kilowatts to megawatts.

Kaufman, H. R.↗

NASA NTRS2002

Performance characteristics of the deep space 1 flight spare ion thruster long duration test, the first 21,300 hours of operation

A long duration test of the DSl flight spare ion thruster (FT2) is presently being conducted at the Jet Propulsion Laboratory. To, date the thruster has accumulated over 23,500 hours of operation, and 190 kg of Xenon propellant, over 230% of the initial design life. The primary objectives of the test include the processing of 200 kg of Xenon propellant, the identification of unknown failure modes, the characterization and drivers of these failure modes, and to measure performance degradation as the thruster wears. The test is fitted with an extensive array of diagnostics to measure engine wear and performance degradation. To date the most notable erosion processes include severe discharge cathode keeper erosion, accelerator grid erosion, reduction in electrical isolation of the neutralizer assembly, and deposit formation within the neutralizer orifice, reducing margin from plume mode. Over the past 23,500 hours of operation, performance degradation has been minimal, and it is anticipated that the above erosion processes will not preclude the thruster from processing over 200 kg of Xenon.

ion thrusters deep space one↗

NASA NTRS2012

Current Density Measurements of an Annular-Geometry Ion Engine

The concept of the annular-geometry ion engine, or AGI-Engine, has been shown to have many potential benefits when scaling electric propulsion technologies to higher power. However, the necessary asymmetric location of the discharge cathode away from thruster centerline could potentially lead to non-uniformities in the discharge not present in conventional geometry ion thrusters. In an effort to characterize the degree of this potential nonuniformity, a number of current density measurements were taken on a breadboard AGI-Engine. Fourteen button probes were used to measure the ion current density of the discharge along a perforated electrode that replaced the ion optics during conditions of simulated beam extraction. Three Faraday probes spaced apart in the vertical direction were also used in a separate test to interrogate the plume of the AGI-Engine during true beam extraction. It was determined that both the discharge and the plume of the AGI-Engine are highly uniform, with variations under most conditions limited to 10% of the average current density in the discharge and 5% of the average current density in the plume. Beam flatness parameter measured 30 mm from the ion optics ranged from 0.85 0.95, and overall uniformity was shown to generally increase with increasing discharge and beam currents. These measurements indicate that the plasma is highly uniform despite the asymmetric location of the discharge cathode.

Shastry, Rohit↗

NASA NTRS2012

Current Density Measurements of an Annular-Geometry Ion Engine

The concept of the annular-geometry ion engine, or AGI-Engine, has been shown to have many potential benefits when scaling electric propulsion technologies to higher power. However, the necessary asymmetric location of the discharge cathode away from thruster centerline could potentially lead to non-uniformities in the discharge not present in conventional geometry ion thrusters. In an effort to characterize the degree of this potential non-uniformity, a number of current density measurements were taken on a breadboard AGI-Engine. Fourteen button probes were used to measure the ion current density of the discharge along a perforated electrode that replaced the ion optics during conditions of simulated beam extraction. Three Faraday probes spaced apart in the vertical direction were also used in a separate test to interrogate the plume of the AGI-Engine during true beam extraction. It was determined that both the discharge and the plume of the AGI-Engine are highly uniform, with variations under most conditions limited to +/-10% of the average current density in the discharge and +/-5% of the average current density in the plume. Beam flatness parameter measured 30 mm from the ion optics ranged from 0.85 - 0.95, and overall uniformity was shown to generally increase with increasing discharge and beam currents. These measurements indicate that the plasma is highly uniform despite the asymmetric location of the discharge cathode.

Shastry, Rohit↗

NASA NTRS2008

Wear Mechanisms in Electron Sources for Ion Propulsion, 2: Discharge Hollow Cathode

The wear of the keeper electrode in discharge hollow cathodes is a major impediment to the implementation of ion propulsion onboard long-duration space science missions. The development of a predictive theoretical model for hollow cathode keeper life has long been sought, but its realization has been hindered by the complexities associated with the physics of the partially ionized gas and the associated erosion mechanisms in these devices. Thus, although several wear mechanisms have been hypothesized, a quantitative explanation of life test erosion profiles has remained incomplete. A two-dimensional model of the partially ionized gas in a discharge cathode has been developed and applied to understand the mechanisms that drove the erosion of the keeper in two long-duration life tests of a 30-cm ion thruster. An extensive set of comparisons between predictions by the numerical simulations and measurements of the plasma properties and of the erosion patterns is presented. It is found that the near-plume plasma oscillations, predicted by theory and observed by experiment, effectively enhance the resistivity of the plasma as well as the energy of ions striking the keeper.

hollow cathode↗

NASA NTRS2016

Wear Testing of the HERMeS Thruster

The Hall-Effect Rocket with Magnetic Shielding (HERMeS) thruster is being developed and tested at NASA GRC and NASA JPL through support of the Space Technology Mission Directorate (STMD) as primary propulsion for the Asteroid Redirect Robotic Mission (ARRM). This thruster is advancing the state-of-the-art of Hall-effect thrusters and is intended to serve as a precursor to higher power systems for human interplanetary exploration. A 2000-hour wear test has been initiated at NASA GRC with the HERMeS Technology Demonstration Unit One and three of four test segments have been completed totaling 728 h of operation. This is the first test of a NASA-designed magnetically shielded thruster to extend beyond 300 hr of continuous operation. Trends in performance, component wear, thermal design, plume properties, and back-sputtered deposition are discussed for two wear-test segments of 246 h and 360 h. The first incorporated graphite pole covers in an electrical configuration where cathode was electrically connected to thruster body. The second utilized traditional alumina pole covers with the thruster body floating. It was shown that the magnetic shielding in both configurations completely eliminated erosion of the boron nitride discharge channel but resulted in erosion of the inner pole cover. The volumetric erosion rate of the graphite pole covers was roughly 2/3 that of the alumina pole covers and the thruster exhibited slightly better performance. Buildup of back-sputtered carbon on the BN channel at a rate of roughly 1.5 μm/kh is shown to have negligible impact on the performance.

Thruster↗