John Yim

Overview of the Development of the Solar Electric Propulsion Technology Demonstration Mission 12.5-kW Hall Thruster

NASA is developing mission concepts for a solar electric propulsion technology demonstration mission. A number of mission concepts are being evaluated including ambitious missions to near Earth objects. The demonstration of a high-power solar electric propulsion capability is one of the objectives of the candidate missions under consideration. In support of NASAs exploration goals, a number of projects are developing extensible technologies to support NASAs near and long term mission needs. Specifically, the Space Technology Mission Directorate Solar Electric Propulsion Technology Demonstration Mission project is funding the development of a 12.5-kW magnetically shielded Hall thruster system to support future NASA missions. This paper presents the design attributes of the thruster that was collaboratively developed by the NASA Glenn Research Center and the Jet Propulsion Laboratory. The paper provides an overview of the magnetic, plasma, thermal, and structural modeling activities that were carried out in support of the thruster design. The paper also summarizes the results of the functional tests that have been carried out to date. The planned thruster performance, plasma diagnostics (internal and in the plume), thermal, wear, and mechanical tests are outlined.

NASA HERMeS Hall Thruster Electrical Configuration Characterization

The NASA Hall Effect Rocket with Magnetic Shielding (HERMeS) 12.5 kW Technology Demonstration Unit-1 (TDU-1) Hall thruster has been the subject of extensive technology maturation in preparation for development into a flight ready propulsion system. Part of the technology maturation was to test the TDU-1 thruster in several ground based electrical configurations to assess the thruster robustness and suitability to successful in-space operation. The ground based electrical configuration testing has recently been demonstrated as an important step in understanding and assessing how a Hall thruster may operate differently in-space compared to ground based testing, and to determine the best configuration to conduct development and qualification testing. This paper describes the electrical configuration testing of the HERMeS TDU-1 Hall thruster in NASA Glenn Research Centers Vacuum Facility 5. The three electrical configurations examined were 1) thruster body tied to facility ground, 2) thruster floating, and 3) thruster body electrically tied to cathode common. The HERMeS TDU-1 Hall thruster was also configured with two different exit plane boundary conditions, dielectric and conducting, to examine the influence on the electrical configuration characterization.

Performance, Facility Pressure Effects, and Stability Characterization Tests of NASA's Hall Effect Rocket with Magnetic Shielding Thruster

NASAs Hall Effect Rocket with Magnetic Shielding (HERMeS) 12.5 kW Technology Demonstration Unit-1 (TDU-1) has been the subject of extensive technology maturation in preparation for flight system development. Part of the technology maturation effort included experimental evaluation of the TDU-1 thruster with conducting and dielectric front pole cover materials in two different electrical configurations. A graphite front magnetic pole cover thruster configuration with the thruster body electrically tied to cathode, and an alumina front pole cover thruster configuration with the thruster body floating were evaluated. Both configurations were also evaluated at different facility background pressure conditions to evaluate background pressure effects on thruster operation. Performance characterization tests found that higher thruster performance was attained with the graphite front pole cover configuration with the thruster electrically tied to cathode. A total thrust efficiency of 68% and a total specific impulse of 2,820 s was demonstrated at a discharge voltage of 600 V and a discharge power of 12.5 kW. Thruster stability regimes were characterized with respect to the thruster discharge current oscillations and with maps of the discharge current-voltage-magnetic field (IVB). Analysis of TDU-1 discharge current waveforms found that lower normalized discharge current peak-to-peak and root mean square magnitudes were attained when the thruster was electrically floated with alumina front pole covers. Background pressure effects characterization tests indicated that the thruster performance and stability were mostly invariant to changes in the facility background pressure for vacuum chamber pressure below 110-5 Torr-Xe (for thruster flow rates of 20.5 mg/s). Power spectral density analysis of the discharge current waveforms showed that increasing the vacuum chamber background pressure resulted in a higher discharge current dominant breathing mode frequency. Finally, IVB maps of the TDU-1 thruster indicated that the discharge current became more oscillatory with higher discharge current peak-to-peak and RMS values with increased facility background pressure at lower thruster mass flow rates; thruster operation at higher flow rates resulted in less change to the thrusters IVB characteristics with elevated background pressure.

Green Propellant Infusion Mission Plume Impingement Analysis

The Green Propellant Infusion Mission (GPIM) is a Technology Demonstration Mission (TDM) project, sponsored by NASA’s Space Technology Mission Directorate (STMD). The goal of GPIM is to demonstrate the capability of a green propulsion system, specifically, one using the monopropellant, AF-M315E. The GPIM propulsion system will be flown as a payload on a Ball Aerospace BCP-100, a small, standardized spacecraft. The propulsion system will have one 22 N thruster for primary divert manuevers and four 1 N thrusters for attitude control. One of the risks identified for GPIM is potential contamination of sensitive areas of the spacecraft from the effluents in the plumes of AF-M315E thrusters. NASA Glenn Research Center (GRC) is conducting activities to mitigate the AF-M315E plume risk. The plume risk mitigation activities include modeling the plume flow fields of the AF-M315E thrusters, assessing the plume impingement on the BCP-100 spacecraft including the impact on the power generating capabilities of the solar arrays, and conducting ground-based plume measurements on an AF-M315E thruster to correlate the plume modeling with plume data. This paper describes the preliminary results from the first activity, plume modeling and plume impingement analysis. Plume flow fields of the 22 N and 1 N thrusters have been modeled using both the method-of-characteristics based Reacting And Multi-Phase (RAMP2) code and the Hypersonic Aerothermodynamics Particle (HAP) Direct Simulation Monte Carlo (DSMC) code. The density and temperature plume flow fields are presented, as well as species concentration at different locations in the plume. Chamber pressures from 400 psia to 100 psia are examined, simulating the GPIM propulsion system operating in blowdown mode. Both equilibrium and frozen flow assumptions are also investigated. The plume impingement on the BCP-100 spacecraft are evaluated, using both the PLume IMPingement program (PLIMP) and also HAP. The heating rates on the spacecraft surfaces are found to be fairly benign. Species flux on the surfaces are also examined, with hydrogen gas found to be the dominant species in the backflow region. The simulations from both models are planned to be compared to forthcoming plume measurement data, when AFM315E thruster testing is conducted.

Carbon Back Sputter Modeling for Hall Thruster Testing

Lifetime requirements for electric propulsion devices, including Hall Effect thrusters, are continually increasing, driven in part by NASAs inclusion of this technology in its exploration architecture. NASA will demonstrate high-power electric propulsion system on the Solar Electric Propulsion Technology Demonstration Mission (SEP TDM). The Asteroid Redirect Robotic mission is one candidate SEP TDM, which is projected to require tens of thousands of thruster life. As thruster life is increased, for example through the use of improved magnetic field designs, the relative influence of facility effects increases. One such effect is the sputtering and redeposition, or back sputter, of facility materials by the high energy thruster plumes. In support of wear testing for the Hall Effect Rocket with Magnetic Shielding (HERMeS) project, the back sputter from a Hall effect thruster plume has been modeled for the NASA Glenn Research Centers Vacuum Facility 5. The predicted wear at a near-worst case condition of 600 V, 12.5 kW was found to be on the order of 1 micron/kh in a fully carbon-lined chamber. A more detailed numerical Monte Carlo code was also modified to estimate back sputter for a detailed facility and pumping configuration. This code demonstrated similar back sputter rate distributions, but is not yet accurately modeling the magnitudes. The modeling has been benchmarked to recent HERMeS wear testing, using multiple microbalance measurements. These recent measurements have yielded values on the order of 1.5 - 2 micron/kh at 600 V and 12.5 kW.

Green Propellant Infusion Mission Program Development and Technology Maturation

The NASA Space Technology Mission Directorates (STMD) Green Propellant Infusion Mission (GPIM) Technology Demonstration Mission (TDM) is comprised of a cross-cutting team of domestic spacecraft propulsion and storable green propellant technology experts. This TDM is led by Ball Aerospace & Technologies Corp. (BATC), who will use their BCP- 100 spacecraft to carry a propulsion system payload consisting of one 22 N thruster for primary divert (DeltaV) maneuvers and four 1 N thrusters for attitude control, in a flight demonstration of the AF-M315E technology. The GPIM project has technology infusion team members from all three major market sectors: Industry, NASA, and the Department of Defense (DoD). The GPIM project team includes BATC, includes Aerojet Rocketdyne (AR), Air Force Research Laboratory, Aerospace Systems Directorate, Edwards AFB (AFRL), NASA Glenn Research Center (GRC), NASA Kennedy Space Center (KSC), and NASA Goddard Space Flight Center (GSFC). STMD programmatic and technology oversight is provided by NASA Marshall Space Flight Center. The GPIM project shall fly an operational AF-M315E green propulsion subsystem on a Ball-built BCP-100 spacecraft.

Hall Thruster Thermal Modeling and Test Data Correlation

The life of Hall Effect thrusters are primarily limited by plasma erosion and thermal related failures. NASA Glenn Research Center (GRC) in cooperation with the Jet Propulsion Laboratory (JPL) have recently completed development of a Hall thruster with specific emphasis to mitigate these limitations. Extending the operational life of Hall thursters makes them more suitable for some of NASAs longer duration interplanetary missions. This paper documents the thermal model development, refinement and correlation of results with thruster test data. Correlation was achieved by minimizing uncertainties in model input and recognizing the relevant parameters for effective model tuning. Throughout the thruster design phase the model was used to evaluate design options and systematically reduce component temperatures. Hall thrusters are inherently complex assemblies of high temperature components relying on internal conduction and external radiation for heat dispersion and rejection. System solutions are necessary in most cases to fully assess the benefits and/or consequences of any potential design change. Thermal model correlation is critical since thruster operational parameters can push some components/materials beyond their temperature limits. This thruster incorporates a state-of-the-art magnetic shielding system to reduce plasma erosion and to a lesser extend power/heat deposition. Additionally a comprehensive thermal design strategy was employed to reduce temperatures of critical thruster components (primarily the magnet coils and the discharge channel). Long term wear testing is currently underway to assess the effectiveness of these systems and consequently thruster longevity.

Green Propellant Infusion Mission program overview and status

The NASA Space Technology Mission Directorates (STMD) Green Propellant Infusion Mission (GPIM) Technology Demonstration Mission (TDM) is comprised of a cross-cutting team of domestic spacecraft propulsion and green technology experts. The GPIM program has technology infusion-team members from all three major market sectors: Industry, NASA, and the Department of Defense (DoD). This team is led and managed by Ball Aerospace & Technologies Corp. (Ball), and includes Aerojet Rocketdyne (AR), Air Force Research Laboratory, Aerospace Systems Directorate, Edwards AFB (AFRL), NASA Glenn Research Center (GRC), and NASA Kennedy Space Center (KSC). STMD programmatic and technology oversight is provided by NASA Marshall Space Flight Center. The GPIM program shall fly an operational AF-M315E green propulsion subsystem on a Ball-built BCP-100 spacecraft.

Flexible tip guides and intermediate catheters: two center experience and a proposed taxonomy

Background Stable access to target lesions is foundational to endovascular therapy, be it in hemorrhagic or ischemic disease. Continued evolution in access technology has resulted in next generation catheters that afford improved trackability and proximal support. Objective Assess safety and patterns of use at two high volume centers, and conceptualize usage patterns. Materials and methods A retrospective review of 608 cases in which a ‘next generation’ catheter was used during 2008–2010 at Cleveland Clinic (Cleveland, Ohio, USA) and throughout 2009–2010 at Emory University Hospital (Atlanta, Georgia, USA) was conducted, and the cases classified by indication. Catheter placement, distal most location, and related complications were recorded and experience summarized. We also reviewed the differences in the catheters and the rationale for catheter selection, as well as relative costs for each approach. Results 311 Neuron 053, 166 Neuron 070, 36 distal access catheter (DAC) 3.9 F, 61 DAC 4.3 F, and 34 DAC 5.2 F catheters were deployed. Of these, 459 placements were in the anterior circulation, 130 in the posterior circulation, 11 in the external carotid artery, and eight were used intravenously. Complication rates were 9/131 (6.9%) for the DAC catheter group, 16/311 (5.1%) for the Neuron 053 group, and 14/166 (8.4%) for the Neuron 070 group (p=0.37, χ2 test). Conclusions Next generation access catheters possess characteristics that blend qualities of traditional microcatheters and stiff guide catheters. There was no statistically significant difference in complication rates between the various catheter families in this small retrospective review, and the complication rates were similar to historical complication rates.

Operational Status of the International Space Station Plasma Contactor Hollow Cathode Assemblies From

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.