A Survey of Xenon Ion Sputter Yield Data and Fits Relevant to Electric Propulsion Spacecraft Integration

A survey of low energy xenon ion impact sputter yields was conducted to provide a more coherent baseline set of sputter yield data and accompanying fits for electric propulsion integration. Data uncertainties are discussed and different available curve fit formulas are assessed for their general suitability. A Bayesian parameter fitting approach is used with a Markov chain Monte Carlo method to provide estimates for the fitting parameters while characterizing the uncertainties for the resulting yield curves

Development Status of a Three-Dimensional Electron Fluid Model for Hall Thruster Plume Simulations

A 3-D electron fluid model has been developed as a stepping stone to fully describe the electron current flow across magnetic fields inside a vacuum chamber and to provide electron flux to solar arrays for spacecraft surface charging model. A detailed description of the numerical treatment of the electric potential solver, including finite-volume formulation, implementation, and the treatment of boundary conditions, are presented in this paper. Verification tests of the model are presented

Characterization of Vacuum Facility Background Gas Through Simulation and Considerations for Electric Propulsion Ground Testing

The background gas in a vacuum facility for electric propulsion ground testing is examined in detail through a series of cold flow simulations using a direct simulation Monte Carlo (DSMC) code. The focus here is on the background gas itself, its structure and characteristics, rather than assessing its interaction and impact on thruster operation. The background gas, which is often incorrectly characterized as uniform, is found to have a notable velocity within a test facility. The gas velocity has an impact on the proper measurement of pressure and the calculation of ingestion flux to a thruster. There are also considerations for best practices for tests that involve the introduction of supplemental gas flows to artificially increase the background pressure. All of these effects need to be accounted for to properly characterize the operation of electric propulsion thrusters across different ground test vacuum facilities

Recommended Practice for Pressure Measurements and Calculation of Effective Pumping Speeds During Electric Propulsion Testing

The electric propulsion community has been implored to establish and implement a set of universally applicable test standards during the research, development, and qualification of electric propulsion systems. Variability between facility-to-facility and more importantly ground-to-flight performance can result in large margins in application or aversion to mission infusion. Performance measurements and life testing under appropriate conditions can be costly and lengthy. Measurement practices must be consistent, accurate, and repeatable. Additionally, the measurements must be universally transportable across facilities throughout the development, qualification, spacecraft integration, and on-orbit performance. A recommended practice for making pressure measurements, pressure diagnostics, and calculating effective pumping speeds with justification is presented

Operational Status of the International Space Station Plasma Contactor Hollow Cathode Assemblies July 2001 to May 2013

The International Space Station has onboard two Aerojet Rocketdyne developed plasma contactor units that perform the function of charge control. The plasma contactor units contain NASA Glenn Research Center developed hollow cathode assemblies. NASA Glenn Research Center monitors the on-orbit operation of the flight hollow cathode assemblies. As of May 31, 2013, HCA.001-F has been ignited and operated 123 times and has accumulated 8072 hours of operation, whereas, HCA.003-F has been ignited and operated 112 times and has accumulated 9664 hours of operation. Monitored hollow cathode ignition times and anode voltage magnitudes indicate that they continue to operate nominally

Operational Status of the International Space Station Plasma Contactor Hollow Cathode Assemblies from July 2011 to May 2013

The International Space Station has onboard two Aerojet Rocketdyne developed plasma contactor units that perform the function of charge control. The plasma contactor units contain NASA Glenn Research Center developed hollow cathode assemblies. NASA Glenn Research Center monitors the onorbit operation of the flight hollow cathode assemblies. As of May 31, 2013, HCA.001-F has been ignited and operated 123 times and has accumulated 8072 hours of operation, whereas, HCA.003-F has been ignited and operated 112 times and has accumulated 9664 hours of operation. Monitored hollow cathode ignition times and anode voltage magnitudes indicate that they continue to operate nominally

Non-Contact Thermal Characterization of NASA's HERMeS Hall Thruster

The thermal characterization test of NASA's 12.5-kW Hall Effect Rocket with Magnetic Shielding has been completed. This thruster was developed to support a number of potential Solar Electric Propulsion Technology Demonstration Mission concepts, including the Asteroid Redirect Robotic Mission concept. As a part of the preparation for this characterization test, an infrared-based, non-contact thermal imaging system was developed to measure the temperature of various thruster surfaces that are exposed to high voltage or plasma. An in-situ calibration array was incorporated into the setup to improve the accuracy of the temperature measurement. The key design parameters for the calibration array were determined in a separate pilot test. The raw data from the characterization test was analyzed though further work is needed to obtain accurate anode temperatures. Examination of the front pole and discharge channel temperatures showed that the thruster temperature was driven more by discharge voltage than by discharge power. Operation at lower discharge voltages also yielded more uniform temperature distributions than at higher discharge voltages. When operating at high discharge voltage, increasing the magnetic field strength appeared to have made the thermal loading azimuthally more uniform

Structural Design Considerations for a 50 kW-Class Solar Array for NASA's Asteroid Redirect Mission

NASA is planning an Asteroid Redirect Mission (ARM) to take place in the 2020s. To enable this multi-year mission, a 40 kW class solar electric propulsion (SEP) system powered by an advanced 50 kW class solar array will be required. Powered by the SEP module (SEPM), the ARM vehicle will travel to a large near-Earth asteroid, descend to its surface, capture a multi-metric ton (t) asteroid boulder, ascend from the surface and return to the Earth-moon system to ultimately place the ARM vehicle and its captured asteroid boulder into a stable distant orbit. During the years that follow, astronauts flying in the Orion multipurpose crew vehicle (MPCV) will dock with the ARM vehicle and conduct extra-vehicular activity (EVA) operations to explore and sample the asteroid boulder. This paper will review the top structural design considerations to successfully implement this 50 kW class solar array that must meet unprecedented performance levels. These considerations include beyond state-of-the-art metrics for specific mass, specific volume, deployed area, deployed solar array wing (SAW) keep in zone (KIZ), deployed strength and deployed frequency. Analytical and design results are presented that support definition of stowed KIZ and launch restraint interface definition. An offset boom is defined to meet the deployed SAW KIZ. The resulting parametric impact of the offset boom length on spacecraft moment of inertias and deployed SAW quasistatic and dynamic load cases are also presented. Load cases include ARM spacecraft thruster plume impingement, asteroid surface operations and Orion docking operations which drive the required SAW deployed strength and damping. The authors conclude that to support NASA's ARM power needs, an advanced SAW is required with mass performance better than 125 W/kg, stowed volume better than 40 kW/cu m, a deployed area of 200 sq m (100 sq m for each of two SAWs), a deployed SAW offset distance of nominally 3-4 m, a deployed SAW quasistatic strength of nominally 0.1 g in any direction, a deployed loading displacement under 2 m, a deployed fundamental frequency above 0.1 Hz and deployed damping of at least 1%. These parameters must be met on top of challenging mission environments and ground testing requirements unique to the ARM project

Hall Thruster Erosion Prediction Using A Hydrodynamic Plasma Model And Sputtering Simulation

One of the challenges facing further utilization of Hall thrusters for space missions is establishing valid lifetime predictions. A combination of the use of a hydrodynamic code and a molecular dynamics code is presented for Hall thruster erosion modeling. The hydrodynamic model simulates the plasma flow within the acceleration channel of an SPT-100 thruster to calculate the ion flux to the walls. The molecular dynamics model simulates the sputtering of boron nitride, a common Hall thruster wall material, to calculate sputter yields at low ion impact energies. Results from both models are shown as well as a qualitative comparison of sample erosion profiles with experimental data

Hybrid-Particle-In-Cell Simulation of Backsputtered Carbon Transport in the Near-Field Plume of a Hall Thruster

Magnetic shielding has eliminated boron nitride erosion as the life limiting mechanism in a Hall thruster but has resulted in erosion of the front magnetic field pole pieces. Recent experiments show that the erosion of graphite pole covers, which are added to protect the magnetic field pole pieces, causes carbon to redeposit on other surfaces, such as boron nitride discharge channel and cathode keeper surfaces. As a part of the risk-reduction activities for Advanced Electric Propulsion System thruster development, this study models transport of backsputtered carbon from the graphite front pole covers and vacuum facility walls. Fluxes, energy distributions, and redeposition rates of backsputtered carbon on the anode, discharge channel, and graphite cathode keeper surfaces are predicted