Electric Propulsion Technology Application Readiness Research Articles

Numerical Simulations of the Partially Ionized Gas in a 100-A LaB<sub>6</sub> Hollow Cathode

Numerical simulations of a hollow cathode with a lanthanum hexaboride (LaB <sub xmlns:mml="http://www.w3.org/1998/Math/MathML" xmlns:xlink="http://www.w3.org/1999/xlink">6</sub> ) emitter operating at 100 A have been performed using the 2-D Orificed Cathode (OrCa2D) code. Results for a variety of plasma properties are presented and compared with laboratory measurements. The large size of the device permits peak electron number densities in the cathode interior that are lower than those established in the NASA Solar Electric Propulsion Technology Application Readiness (NSTAR) hollow cathode, which operates at a 7.3× lower discharge current and 3.2× lower mass flow rate. The maximum electron current density also is lower in the LaB <sub xmlns:mml="http://www.w3.org/1998/Math/MathML" xmlns:xlink="http://www.w3.org/1999/xlink">6</sub> cathode, by 4.2×, due to the larger orifice size. Simulations and direct measurements show that at 12 sccm of xenon flow the peak emitter temperature is in the range of 1630°C-1666°C. It is also found that the conditions for the excitement of current-driven streaming instabilities and ion-acoustic turbulence (IAT) are satisfied in this cathode, similarly to what was found in the past in its smaller counterparts like the NSTAR cathode. Based on numerical simulations, it has long been argued that these instabilities may be responsible for the anomalously large ion energies that have been measured in these discharges as well as for the enhancement of the plasma resistivity. Direct measurements of the turbulent spectra and confirmation of the presence of IAT in this cathode have now been completed. Interpolation of the measured anomalous collision frequency based on slightly different operating conditions than the one in the numerical simulations suggests good agreement with the computed values.

Particle-Based Plasma Simulations for an Ion EngineDischarge Chamber

A particle-based model with a Monte Carlo collision model has been developed to study the plasma inside the discharge chamber of an ion engine. This model tracks five major particle types inside the discharge chamber in detail: xenon neutrals, singly charged xenon ions, doubly charged xenon ions, secondary electrons, and primary electrons. Both electric and magnetic field effects are included in the calculation of the charged particle's motion. The electric fields inside the discharge chamber are computed using a new approach. Also, detailed particle collision mechanisms are enabled. Validation of this computational model has been made on NASA's three-ring Solar Electric Propulsion Technology Application Readiness Program discharge chamber, at the 2.29 kW input power, 1.76 A beam current, and 1100 V beam voltage operating condition. Comparisons of numerical simulation results with experimental measurements are found to have good agreement. The computed ion beam current differs from experiments by 1% and the computed discharge current differs from experiments by 22%. The plasma ion production cost compares within 7% and the beam ion production cost compares within 16% of the experimental values. The overall computed thruster efficiency is found to differ from experiments by 11 %. In addition, steady-state results are given for particle number density distributions, kinetic energy, particle energy loss mechanisms, and current density collected at the chamber walls.

Langmuir Probe Studies of Magnetic Confinement in an Ion Thruster Discharge Plasma

Magnetic confinement studies on a 30-cm-diameter ion thruster were performed to determine the dependence of plasma confinement and uniformity on the ring-cusp magnetic field. Four primary cases were investigated to determine the effects of an additional magnetic cusp, increasing the strength of the highest value closed magnetic contour line, and varying the magnetic field free volume. A laboratory model NASA Solar Electric Propulsion Technology Application Readiness engine was modified to investigate three- and four-ring-cusp geometries. Electrical parameters and langmuir probe sweeps in the discharge chamber were used to measure the performance of each configuration. Increasing the strength of the closed magnetic contour line reduces ion loss to the anode, resulting in a reduction in discharge power for a given beam current. Similarly, increasing the magnetic field free volume in the near-grid region improves plasma uniformity, removing the on-axis current density peak responsible for accelerator grid erosion. The enhanced magnetic circuit geometries investigated resulted in a 20% reduction in discharge loss and discharge current at the TH15 (2.3 kW) throttle point, with similar gains over the full throttle range. The reduction in discharge power and peak current density can significantly increase the total throughout per engine by limiting wear mechanisms that limit thruster life.

Thermal Model of the Hollow Cathode Using Numerically Simulated Plasma Fluxes

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.

Numerical Thermal Model of NASA Solar Electric Propulsion Technology Application Readiness Ion Thruster

A thermal computer model of the 30-cm NASA solar electric propulsion technology application readiness (NSTAR) xenon ion thruster has been produced using a lumped-parameter thermal nodal-network scheme. This model contains 104 nodes on the thruster and was implemented using SINDA and TRASYS on various UNIX workstations. The model includes the radiation and conduction heat transfer, the effect of plasma interaction on the thruster, andan account for Ž nely perforated surfaces. Themodelwasdeveloped in conjunctionwith anNSTAR thruster outŽ tted with approximately 20 thermocouples for thermal testing at the JohnH. Glenn Research Center. The results of these experiments were used to calibrate and conŽ rm the computer model Ž rst without and then with the plasma interaction. The calibrated model was able to predict discharge chamber temperatures to within 10±C of measured temperatures. To demonstrate the ability of the model under various circumstances, the heat  ux was examined for a thruster operating in a deep-space environment.