Temperature Distributions in Hollow Cathode Emitters

Polk, James E.; Marrese, Colleen M.; Dang, Lisa; Johnson, L. K.; Thornber, Ben

doi:10.2514/6.2004-4116

Cited by 38 publications

(23 citation statements)

References 9 publications

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“…14,15,16,23 Cathode-to-ground voltage results in Figure 8 indicate that, in general, the cathode-toground voltage became more negative as the flow rate was increased. At a flow rate of 19 sccm, configuration 1 operated in plume mode for all discharge currents.…”

Section: Resultsmentioning

confidence: 90%

Development and Testing of High Current Hollow Cathodes for High Power Hall Thrusters

Kamhawi

Noord²

2012

48th AIAA/ASME/SAE/ASEE Joint Propulsion Conference &Amp;amp; Exhibit

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NASA's Office of the Chief Technologist In-Space Propulsion project is sponsoring the testing and development of high power Hall thrusters for implementation in NASA missions. As part of the project, NASA Glenn Research Center is developing and testing new high current hollow cathode assemblies that can meet and exceed the required discharge current and life-time requirements of high power Hall thrusters. This paper presents test results of three high current hollow cathode configurations. Test results indicated that two novel emitter configurations were able to attain lower peak emitter temperatures compared to state-of-the-art emitter configurations. One hollow cathode configuration attained a cathode orifice plate tip temperature of 1132 °C at a discharge current of 100 A. More specifically, test and analysis results indicated that a novel emitter configuration had minimal temperature gradient along its length. Future work will include cathode wear tests, and internal emitter temperature and plasma properties measurements along with detailed physics based modeling. NomenclatureA = Ampere A o = Universal constant, 120 A/cm 2 C = fit coefficient J d = discharge current density, A/cm 2 J e = return electron current density, A/cm 2 J em = emission cathode current density, A/cm 2 J ion = return ion current density, A/cm 2 J orif = orifice current density, A/cm 2 k = Boltzmann constant, 1.38×10 -23 Joules/K m e = electron mass, 9.1x10 -31 kg n e = electron number density, m -3 q e = electron charge, 1.609×10 -19 Coulomb sccm = standard cubic centimeter per minute T e = electron temperature, eV T emit = emitter temperature, K t = lifetime, hours t depth = time to barium depletion depth, hours V a = energy of activation x = axial position y depth = depth into emitter α = emitter constant, eV/K φ w = emitter work function, eV φ o = temperature independent work function, eV sheath = sheath potential, V * Aerospace Engineer, Propulsion and Propellants Branch, hani.kamhawi-1@nasa.gov, Associate AIAA Fellow. † Aerospace Engineer, Propulsion and Propellants Branch, jonathan.l.vannoord@nasa.gov, Senior AIAA Member.

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Section: Resultsmentioning

confidence: 90%

Development and Testing of High Current Hollow Cathodes for High Power Hall Thrusters

Kamhawi

Noord²

2012

48th AIAA/ASME/SAE/ASEE Joint Propulsion Conference &Amp;amp; Exhibit

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“…The system was calibrated in situ using the hollow cathode heater and thermocouples on the orifice plate, which compensates for cavity and surface effects on the effective emittance. Extensive testing demonstrated that the measurements were repeatable at a given operating point to within 1-2 C. In regards to the thermocouples, under nearly isothermal conditions during calibration, the readings agreed to within 2 C. During cathode operation, however, they differed by 10-15 C. A more detailed description of the system and methods used for the measurements is provided in [19,20]. Results from an earlier test with a different insert and slightly lower flow rate are also plotted and show similar behavior.…”

Section: Resultsmentioning

confidence: 95%

Thermal Model of the Hollow Cathode Using Numerically Simulated Plasma Fluxes

Katz

Mikellides

Polk

et al. 2007

Journal of Propulsion and Power

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The first results from a hollow cathode thermal model are presented. The thermal model uses the spatially distributed plasma fluxes, calculated by a two-dimensional axisymmetric model of the plasma inside the emitter region, as the heat source to predict the hollow cathode and insert temperatures. The computed insert temperature profiles are compared with measured values for a cathode similar to the 0.635-cm diameter discharge cathode used in the NASA solar electric propulsion technology applications readiness ion engine. The insert temperatures can be used to predict cathode life from barium depletion. The present results and related experimental studies yield three important conclusions. First, the emitter in the NASA solar electric propulsion technology applications readiness hollow cathode does not operate in the emission-limited regime. The thermionic electron current is about 20 A higher than the discharge current and requires significant reverse electron flux from the plasma to satisfy current continuity. Second, the high-plasma density near the centerline of the cathode results in power deposition on the orifice plate that is more than twice the emitter power deposition. Third, despite a higher heat load to the orifice plate, the operating temperature where it would be measured by a thermocouple is approximately 100 C lower than the emitter. The lower orifice plate temperature is due to poor thermal contact between the emitter and the cathode tube and higher than anticipated radiative losses from the external surface of the heater. Nomenclature F = geometrical configuration factor IP = ionization potential, V j emission = emission current density, A=m 2 j the e = plasma thermal current density, A=m 2 j i = plasma ion current density, A=m 2 _ q rad = heat loss by radiation, W=m 2 _ q emission = heat loss by electron emission, W=m 2 _ q e = heating by plasma electrons, W=m 2 _ q i = heating by plasma ion flux, W=m 2 T = material temperature, K T e = plasma electron temperature, eV WF = emitter work function, eV ij = Kronecker delta " = material emissivity = material thermal conductivity, W=m K = Stephan-Boltzman constant, W=m 2 K 4 sheath = sheath potential, V Introduction H IGH-POWER, long-life electric thrusters were considered necessary as NASA planned many of the challenging missions under the Prometheus program. These missions might have required an increase in both life and emission current compared with hollow cathodes used on NASA's Deep Space 1 spacecraft. Although both the discharge and neutralizer hollow cathodes were operating at the end of the NASA solar electric propulsion technology applications readiness (NSTAR) thruster 30,000-h extended life test [1], the future missions may be even more challenging. A previous life test of the International Space Station plasma contactor hollow cathode [2] ended after 28,000 h, when the cathode would no longer ignite. One scenario for end of life is when the insert can no longer provide free barium to lower the work function and the remaining barium in th...

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“…This corresponds to an orifice plate temperature of at most 1030 °C, as shown in Figure 1. It has been shown that the emitter on a cathode can be up to 90 °C cooler 26 or 100 °C hotter 27 than the orifice plate temperature. For the purposes of this study, the lifetime results are shown over a +/-100 °C temperature span on the orifice plate temperatures.…”

Section: Insert Barium Depletionmentioning

confidence: 99%

“…25 The peak insert temperature measured on a similar cathode was 1180 °C at 12 A of emission current. 26 Using that peak temperature and an estimated depletion depth of 500 µm, the model outlined in Refs. 21 and 23 yields a lifetime estimate of 30 khrs.…”

Section: Insert Barium Depletionmentioning

confidence: 99%

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Lifetime Assessment of the NEXT Ion Thruster

VanNoord

2007

43rd AIAA/ASME/SAE/ASEE Joint Propulsion Conference &Amp;amp; Exhibit

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Ion thrusters are low thrust, high specific impulse devices with required operational lifetimes on the order of 10,000-100,000 hours. The NEXT ion thruster is the latest generation of ion thrusters under development. The NEXT ion thruster currently has a qualification level propellant throughput requirement of 450 kg of xenon, which corresponds to roughly 22,000 hours of operation at the highest throttling point. Currently, a NEXT engineering model ion thruster with prototype model ion optics is undergoing a long duration test to determine wear characteristics and establish propellant throughput capability. The NEXT thruster includes many improvements over previous generations of ion thrusters, but two of its component improvements have a larger effect on thruster lifetime. These include the ion optics with tighter tolerances, a masked region and better gap control, and the discharge cathode keeper material change to graphite. Data from the NEXT 2000 hour wear test, the NEXT long duration test, and further analysis is used to determine the expected lifetime of the NEXT ion thruster. This paper will review the predictions for all of the anticipated failure mechanisms. The mechanisms will include wear of the ion optics and cathode's orifice plate and keeper from the plasma, depletion of low work function material in each cathode's insert, and spalling of material in the discharge chamber leading to arcing. Based on the analysis of the NEXT ion thruster, the first failure mode for operation above a specific impulse of 2000 s is expected to be the structural failure of the ion optics at 750 kg of propellant throughput, 1.7 times the qualification requirement. An assessment based on mission analyses for operation below a specific impulse of 2000 s indicates that the NEXT thruster is capable of double the propellant throughput required by these missions.

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Temperature Distributions in Hollow Cathode Emitters

Cited by 38 publications

References 9 publications

Development and Testing of High Current Hollow Cathodes for High Power Hall Thrusters

Development and Testing of High Current Hollow Cathodes for High Power Hall Thrusters

Thermal Model of the Hollow Cathode Using Numerically Simulated Plasma Fluxes

Lifetime Assessment of the NEXT Ion Thruster

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