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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.

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NASA NTRS2018

Update on the Nested Hall Thruster Subsystem for the NextSTEP XR-100 Program

Under the NextSTEP program led by Aerojet Rocketdyne in collaboration with NASA Glenn Research Center the University of Michigan, and the Jet Propulsion Laboratory, the XR-100 100 kW Electric Propulsion system is being developed to Technology Readiness Level (TRL) 5. As part of this program, the X3, a Nested Hall Thruster (NHT) designed to operate at powers up to 200kW, is being further developed through parallel modeling and experimental efforts with the ultimate goal of supporting a 100kW-100hr system test in the final year of the NextSTEP program. Recent developments for the X3 subsystem are presented including a summary of testing and modeling results and design updates in anticipation of a risk reduction test scheduled for the summer of 2018.

Jorns, Benjamin A.↗
NASA NTRS2015

Propulsion Technology Assessment: Science and Enabling Technologies to Explore the Interstellar Medium

The Advanced Concepts Office at NASA’s George C. Marshall Space Flight Center conducted a study to assess what low-thrust advanced propulsion system candidates, existing and near term, could deliver a small, Voyager-like satellite to our solar system’s heliopause, approximately 100 AU from the sun, within 10 years. The advanced propulsion system trade study consisted of three candidates, including a Magnetically Shielded Miniature Hall thruster, a solar sail and an electric sail. A second analysis was conducted to determine which solid rocket motor kick stage(s) would be required to provide additional thrust at various points in the trajectory, assuming a characteristic energy capability provided by a Space Launch System Block 1B vehicle architecture carrying an 8.4 meter payload fairing. Two trajectory profiles were considered, including an escape trajectory using a Jupiter gravity assist and an escape trajectory first performing a Jupiter gravity assist followed by an Oberth maneuver around the sun and an optional Saturn gravity assist. Results indicated that if the Technology Readiness Level of an electric sail could be increased in time, this technology could not only enable a satellite to reach 100 AU in 10 years but it could potentially do so in even less time.

Hopkins, Randall C.↗
NASA NTRS1997

Over Powering Solar System Exploration

This paper describes the new mission possibilities and technology requirements for a high power solar Electric Propulsion (SEP) system using a Taurus-class launch vehicle. A 10-kW Hall effect Thruster with Anode Layer (TAL) running direct-drive off a high voltage solar array could produce enough thrust to enable very rapid missions with a small sciencecraft to various destinations throughout the solar system.

Solar↗
NASA NTRS2012

Experimental Investigation of a Direct-drive Hall Thruster and Solar Array System at Power Levels up to 10 kW

As NASA considers future exploration missions, high-power solar-electric propulsion (SEP) plays a prominent role in achieving many mission goals. Studies of high-power SEP systems (i.e. tens to hundreds of kilowatts) suggest that significant mass savings may be realized by implementing a direct-drive power system, so NASA recently established the National Direct-Drive Testbed to examine technical issues identified by previous investigations. The testbed includes a 12-kW solar array and power control station designed to power single and multiple Hall thrusters over a wide range of voltages and currents. In this paper, single Hall thruster operation directly from solar array output at discharge voltages of 200 to 450 V and discharge powers of 1 to 10 kW is reported. Hall thruster control and operation is shown to be simple and no different than for operation on conventional power supplies. Thruster and power system electrical oscillations were investigated over a large range of operating conditions and with different filter capacitances. Thruster oscillations were the same as for conventional power supplies, did not adversely affect solar array operation, and were independent of filter capacitance from 8 to 80 ?F. Solar array current and voltage oscillations were very small compared to their mean values and showed a modest dependence on capacitor size. No instabilities or anomalous behavior were observed in the thruster or power system at any operating condition investigated, including near and at the array peak power point. Thruster startup using the anode propellant flow as the power 'switch' was shown to be simple and reliable with system transients mitigated by the proper selection of filter capacitance size. Shutdown via cutoff of propellant flow was also demonstrated. A simple electrical circuit model was developed and is shown to have good agreement with the experimental data.

Snyder, John S.↗
NASA NTRS2024

SUbsonic Single Aft eNgine (SUSAN) Power/Propulsion System Control Architecture Updates

The development of hybrid-electric propulsion technology for commercial transport aircraft presents opportunities for new designs that can reduce fuel consumption and greenhouse gas emissions and improve safety and reliability compared to modern aircraft. The SUbsonic Single Aft eNgine (SUSAN) Electrofan is a conceptual design for a transport aircraft with a series/parallel partial hybrid-electric propulsion system being developed by NASA as a reference design for a regional transport aircraft with a high degree of electrification. The SUSAN concept incorporates multiple tightly coupled power, propulsion, and flight control systems that introduce new challenges to the control design process, requiring a hierarchical and more coordinated control approach. This report summarizes updates to the SUSAN power/propulsion system (PPS) model and control architecture made in preparation for planned flight simulator and hardware-in-the-loop testing at the NASA Glenn Research Center. These updates include a new electrical power system model developed using the NASA-developed Electrical Modeling and Thermal Analysis Toolbox (EMTAT) and a new PPS control architecture that improves the operational flexibility and responsiveness of the propulsion system. The performance of the updated controller is demonstrated and discussed.

SUSAN↗
NASA NTRS2011

Power and Propulsion System Design for Near-Earth Object Robotic Exploration

Near-Earth Objects (NEOs) are exciting targets for exploration; they are relatively easy to reach but relatively little is known about them. With solar electric propulsion, a vast number of interesting NEOs can be reached within a few years and with extensive flexibility in launch date. An additional advantage of electric propulsion for these missions is that a spacecraft can be small, enabling a fleet of explorers launched on a single vehicle or as secondary payloads. Commercial, flight-proven Hall thruster systems have great appeal based on their performance and low cost risk, but one issue with these systems is that the power processing units (PPUs) are designed for regulated spacecraft power architectures which are not attractive for small NEO missions. In this study we consider the integrated design of power and propulsion systems that utilize the capabilities of existing PPUs in an unregulated power architecture. Models for solar array and engine performance are combined with low-thrust trajectory analyses to bound spacecraft design parameters for a large class of NEO missions, then detailed array performance models are used to examine the array output voltage and current over a bounded mission set. Operational relationships between the power and electric propulsion systems are discussed, and it is shown that both the SPT-100 and BPT-4000 PPUs can perform missions over a solar range of 0.7 AU to 1.5 AU - encompassing NEOs, Venus, and Mars - within their operable input voltage ranges. A number of design trades to control the array voltage are available, including cell string layout, array offpointing during mission operations, and power draw by the Hall thruster system.

Near Earth Object (NEO)↗
NASA NTRS2013

Fairfield Plume Measurement and Analysis on the NASA-300M and NASA-300MS

NASA is developing a 10- to 15-kW Hall thruster system to support future NASA missions. This activity is funded under the Space Technology Mission Directorate Solar Electric Propulsion Technology Demonstration Mission project. As a part of the development process, the NASA-300M, a 20-kW Hall thruster, was modified to incorporate the magnetic shielding concept and named the NASA-300MS. This activity was undertaken to assess the viability of using the magnetic shielding concept on a high-power Hall thruster to greatly reduce discharge channel erosion. This paper reports on the study to characterize the far-field plumes of the NASA-300M and NASA-300MS. Diagnostics deployed included a polarlyswept Faraday probe, a Wien filter (ExB probe), a retarding potential analyzer, and a Langmuir probe. During the study, a new, more accurate, integration method for analyzing Wien filter probe data was implemented and effect of secondary electron emission on the Faraday probe data was treated. Comparison of the diagnostic results from the two thrusters showed that the magnetically shielded version performed with ~2 percent higher voltage utilization efficiency, ~2 percent lower plume divergence efficiency, and ~2 percent lower mass utilization efficiency compared to the baseline version. The net change in efficiency is within the aggregate measurement uncertainty so the overall performance is roughly equal for the two versions of the thruster. Anode efficiency calculated from thrust stand measurement corroborates this finding.

Hall thruster↗
NASA NTRS2003

Direct Drive Hall Thruster System Development

The sta:us of development of a Direct Drive Ha!! Thruster System is presented. 13 the first part. a s:udy of the impacts to spacecraft systems and mass benefits of a direct-drive architecture is reviewed. The study initially examines four cases of SPT-100 and BPT-4000 Hall thrusters used for north-south station keeping on an EXPRESS-like geosynchronous spacecraft and for primary propulsion for a Deep Space- 1 based science spacecraft. The study is also extended the impact of direct drive on orbit raising for higher power geosynchronous spacecraft and on other deep space missions as a function of power and delta velocity. The major system considerations for accommodating a direct drive Hall thruster are discussed, including array regulation, system grounding, distribution of power to the spacecraft bus, and interactions between current-voltage characteristics for the arrays and thrusters. The mass benefit analysis shows that, for the initial cases, up to 42 kg of dry mass savings is attributable directly to changes in the propulsion hardware. When projected mass impacts of operating the arrays and the electric power system at 300V are included, up to 63 kg is saved for the four initial cases. Adoption of high voltage lithium ion battery technology is projected to further improve these savings. Orbit raising of higher powered geosynchronous spacecraft, is the mission for which direct drive provides the most benefit, allowing higher efficiency electric orbit raising to be accomplished in a limited period of time, as well as nearly eliminating significant power processing heat rejection mass. The total increase in useful payload to orbit ranges up to 278 kg for a 25 kW spacecraft, launched from an Atlas IIA. For deep space missions, direct drive is found to be most applicable to higher power missions with delta velocities up to several km/s , typical of several Discovery-class missions. In the second part, the status of development of direct drive propulsion power electronics is presented. The core of this hardware is the heater-keeper-magnet supply being qualified for the BPT-4000 by Aerojet. A breadboard propulsion power unit is in fabrication and is scheduled for delivery late in 2003.

Hoskins, W. Andrew↗
NASA NTRS2011

MW-Class Electric Propulsion System Designs

Electric propulsion systems are well developed and have been in commercial use for several years. Ion and Hall thrusters have propelled robotic spacecraft to encounters with asteroids, the Moon, and minor planetary bodies within the solar system, while higher power systems are being considered to support even more demanding future space science and exploration missions. Such missions may include orbit raising and station-keeping for large platforms, robotic and human missions to near earth asteroids, cargo transport for sustained lunar or Mars exploration, and at very high-power, fast piloted missions to Mars and the outer planets. The Advanced In-Space Propulsion Project, High Efficiency Space Power Systems Project, and High Power Electric Propulsion Demonstration Project were established within the NASA Exploration Technology Development and Demonstration Program to develop and advance the fundamental technologies required for these long-range, future exploration missions. Under the auspices of the High Efficiency Space Power Systems Project, and supported by the Advanced In-Space Propulsion and High Power Electric Propulsion Projects, the COMPASS design team at the NASA Glenn Research Center performed multiple parametric design analyses to determine solar and nuclear electric power technology requirements for representative 300-kW class and pulsed and steady-state MW-class electric propulsion systems. This paper describes the results of the MW-class electric power and propulsion design analysis. Starting with the representative MW-class vehicle configurations, and using design reference missions bounded by launch dates, several power system technology improvements were introduced into the parametric COMPASS simulations to determine the potential system level benefits such technologies might provide. Those technologies providing quantitative system level benefits were then assessed for technical feasibility, cost, and time to develop. Key assumptions and primary results of the COMPASS MW-class electric propulsion power system study are reported, and discussion is provided on how the analysis might be used to guide future technology investments as NASA moves to more capable high power in-space propulsion systems.

LaPointe, Michael R.↗
NASA NTRS2017

13kW Advanced Electric Propulsion Flight System Development and Qualification

The next phase of robotic and human deep space exploration missions is enhanced by high performance, high power solar electric propulsion systems for large-scale science missions and cargo transportation. Aerojet Rocketdynes Advanced Electric Propulsion System (AEPS) program is completing development, qualification and delivery of five flight 13.3kW EP systems to NASA. The flight AEPS includes a magnetically-shielded, long-life Hall thruster, power processing unit (PPU), xenon flow controller (XFC), and intrasystem harnesses. The Hall thruster, originally developed and demonstrated by NASAs Glenn Research Center and the Jet Propulsion Laboratory, operates at input powers up to 12.5kW while providing a specific impulse over 2600s at an input voltage of 600V. The power processor is designed to accommodate an input voltage range of 95 to 140V, consistent with operation beyond the orbit of Mars. The integrated system is continuously throttleable between 3 and 13.3kW. The program has completed the system requirement review; the system, thruster, PPU and XFC preliminary design reviews; development of engineering models, and initial system integration testing. This paper will present the high power AEPS capabilities, overall program and design status and the latest test results for the 13.3kW flight system development and qualification program.

Jackson, Jerry↗
NASA NTRS2019

Plasma simulations for the assessment of pole erosion in the Magnetically Shielded Miniature Hall Thruster (MaSMi)

Numerical simulations have been conducted to assess the erosion rates at the pole covers of the magnetically shielded miniature (MaSMi) Hall thruster. MaSMi is part of the Ascendant Sub-kW Transcelestial Electric Propulsion System (ASTRAEUS), whose design requirements demand a xenon throughput of 100 kg, thus enabling deep space missions using SmallSats. The results of our simulations show that the proposed thickness of the graphite pole covers will not be completely eroded at end of life. We find that the erosion rates are larger near the corners of the covers, which are directly exposed to the plasma in the acceleration channel.

Chaplin, Vernon H.↗
NASA NTRS2019

The Application of Advanced Electric Propulsion on the NASA Power and Propulsion Element (PPE)

NASA is charged with landing the first American woman and next American man on the South Pole of the Moon by 2024. To meet this challenge, NASA's Gateway will develop and deploy critical infrastructure required for operations on the lunar surface and that enables a sustained presence on and around the moon. NASA's Power and Propulsion Element (PPE), the first planned element of NASA's cis-lunar Gateway, leverages prior and ongoing NASA and U.S. industry investments in high-power, long-life solar electric propulsion technology investments. NASA awarded a PPE contract to Maxar Technologies to demonstrate a 2,500 kg xenon capacity, 50 kW-class SEP spacecraft that meets Gateway's needs, aligns with industry's heritage spacecraft buses, and allows extensibility for NASA's Mars exploration goals. Maxar's PPE concept design, is based directly on their high heritage, modular, highly reliable 1300-series bus architecture. The electric propulsion system features two 13 kW Advanced Electric Propulsion (AEPS) strings from Aerojet Rocketdyne and a Maxar-developed system comprised of four Busek 6 kW Hall-effect thrusters mounted in pairs on large range of motion pointing arms with four 6 kW-class, SPT-140-based PPUs. NASA is continuing to develop the 13 kW AEPS system through a contract with Aerojet Rocketdyne. In addition to the flight demonstration of an advanced electric propulsion system on PPE, a government-furnished plasma diagnostics package is planned to assess on-orbit performance characteristics and vehicle interactions. The paper will present overviews of NASA's Gateway and the PPE Project, the Maxar ion propulsion subsystem, the status of the two electric propulsion system developments, and the implementation of the plasma diagnostics package on the Maxar PPE spacecraft. The project is currently heading into SRR, with the propulsion build scheduled for 2021, and launch in 2022.

Herman, Daniel A.↗
NASA NTRS2016

Facility Effect Characterization Test of NASA's HERMeS Hall Thruster

A test to characterize the effect of varying background pressure on NASA's 12.5-kW Hall Effect Rocket with Magnetic Shielding had being completed. This thruster is the baseline propulsion system for the Solar Electric Propulsion Technology Demonstration Mission (SEP TDM). Potential differences in thruster performance and oscillation characteristics when in ground facilities versus on-orbit are considered a primary risk for the propulsion system of the Asteroid Redirect Robotic Mission, which is a candidate for SEP TDM. The first primary objective of this test was to demonstrate that the tools being developed to predict the zero-background-pressure behavior of the thruster can provide self-consistent results. The second primary objective of this test was to provide data for refining a physics-based model of the thruster plume that will be used in spacecraft interaction studies. Diagnostics deployed included a thrust stand, Faraday probe, Langmuir probe, retarding potential analyzer, Wien filter spectrometer, and high-speed camera. From the data, a physics-based plume model was refined. Comparisons of empirical data to modeling results are shown.

Electric Propulsion↗
NASA NTRS2015

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 Center's 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.

HERMeS↗
NASA NTRS2016

NASA HERMeS Hall Thruster Electrical Configuration Characterization

NASAs 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 presentation will cover the electrical configuration testing of the TDU-1 HERMeS Hall thruster in NASA Glenn Research Centers Vacuum Facility 5. The three electrical configurations examined are the thruster body tied to facility ground, thruster floating, and finally the thruster body electrically tied to cathode common. The TDU-1 HERMeS was configured with two different exit plane boundary conditions, dielectric and conducting, to examine the influence on the electrical configuration characterization.

Facility↗
NASA NTRS2021

Electric Propulsion for the Psyche Mission: One Year to Launch

NASA’s Psyche mission will launch in 2022 and begin a 3.6-year cruise to the metallic asteroid Psyche, the largest metal asteroid in the solar system. The baseline spacecraft design is a hybrid of JPL’s deep-space heritage subsystems with commercial partner Maxar’s electric propulsion, power, and structure subsystems. The Psyche mission will feature the first use of Hall thrusters for a NASA mission, and the first use of Hall thrusters beyond cis-lunar space. This paper will provide an overview of the electric propulsion subsystem and details of the flight system implementation for the Psyche mission. The concept of operations is discussed including subsystem-level operations, the initial checkout after launch, and cruise and orbital operations. Command and control interfaces with the spacecraft and fault protection design are presented. Finally, the status of subsystem assembly and test is described.

Li, Julie↗
NASA NTRS2007

Throttling Impacts on Hall Thruster Performance, Erosion, and Qualification for NASA Science Missions

With the SMART-1, Department of Defense, and commercial industry successes in Hall thruster technologies, NASA has started considering Hall thrusters for science missions. The recent Discovery proposals included a Hall thruster science mission and the In-Space Propulsion Project is investing in Hall thruster technologies. As the confidence in Hall thrusters improve, ambitious multi-thruster missions are being considered. Science missions often require large throttling ranges due to the 1/r(sup 2) power drop-off from the sun. Deep throttling of Hall thrusters will impact the overall system performance. Also, Hall thrusters can be throttled with both current and voltage, impacting erosion rates and performance. Last, electric propulsion thruster lifetime qualification has previously been conducted with long duration full power tests. Full power tests may not be appropriate for NASA science missions, and a combination of lifetime testing at various power levels with sufficient analysis is recommended. Analyses of various science missions and throttling schemes using the Aerojet BPT-4000 and NASA 103M HiVHAC thruster are presented.

Dankanich, John W.↗
NASA NTRS2006

Evaluation of High-Power Solar Electric Propulsion using Advanced Ion, Hall, MPD, and PIT Thrusters for Lunar and Mars Cargo Missions

This paper presents the results of mission analyses that expose the advantages and disadvantages of high-power (MWe-class) Solar Electric Propulsion (SEP) for Lunar and Mars Cargo missions that would support human exploration of the Moon and Mars. In these analyses, we consider SEP systems using advanced Ion thrusters (the Xenon [Xe] propellant Herakles), Hall thrusters (the Bismuth [Bi] propellant Very High Isp Thruster with Anode Layer [VHITAL], magnetoplasmadynamic (MPD) thrusters (the Lithium [Li] propellant Advanced Lithium-Fed, Applied-field Lorentz Force Accelerator (ALFA2), and pulsed inductive thruster (PIT) (the Ammonia [NH3] propellant Nuclear-PIT [NuPIT]). The analyses include comparison of the advanced-technology propulsion systems (VHITAL, ALFA2, and NuPIT) relative to state-of-theart Ion (Herakles) propulsion systems and quantify the unique benefits of the various technology options such as high power-per-thruster (and/or high power-per-thruster packaging volume), high specific impulse (Isp), high-efficiency, and tankage mass (e.g., low tankage mass due to the high density of bismuth propellant). This work is based on similar analyses for Nuclear Electric Propulsion (NEP) systems.

Lunar cargo↗