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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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At least 343 records · Page 19
All sourcesNASA NTRSDOE OSTI
NASA NTRS1989

Computer Control For Ion Engines

Computer system controls start-up, steady-state operation, throttling, and shutdown of pair of xenon-ion propulsion engines. Controls direction of thrust of each engine through operation of gimbal stepping motors. Controls valves in propellant-storage and propellant-distribution system. Control software establishes and maintains efficient, stable operation over entire range of operating variables, and throttles engines to any point within range. Computer operates power supplies, valves, and flow controllers of two ion engines, ion-neutralizer subsystem, and other equipment. Designed for use in interplanetary flight, system adaptable to industrial use in ion-beam deposition of thin films.

Brophy, John R.↗
NASA NTRS2017

Overview of the Development and Mission Application of the Advanced Electric Propulsion System (AEPS)

NASA remains committed to the development and demonstration of a high-power solar electric propulsion capability for the Agency. NASA is continuing to develop the 14 kilowatt Advanced Electric Propulsion System (AEPS), which has recently completed an Early Integrated System Test and System Preliminary Design Review. NASA continues to pursue Solar Electric Propulsion (SEP) Technology Demonstration Mission partners and mature high-power SEP mission concepts. The recent announcement of the development of a Power and Propulsion Element (PPE) as the first element of an evolvable human architecture to Mars has replaced the Asteroid Redirect Robotic Mission as the most probable first application of the AEPS Hall thruster system. This high-power SEP capability, or an extensible derivative of it, has been identified as a critical part of an affordable, beyond-low-Earth-orbit, manned-exploration architecture. This paper presents the status of the combined NASA and Aerojet AEPS development activities and updated mission concept for implementation of the AEPS hardware as part of the ion propulsion system for a PPE.

ion engines↗
NASA NTRS1997

Development of Power Electronics for a 0.2kW-Class Ion Thruster

Applications that might benefit from low power ion propulsion systems include Earth-orbit magnetospheric mapping satellite constellations, low Earth-orbit satellites, geosynchronous Earth-orbit satellite north-south stationkeeping, and asteroid orbiters. These spacecraft are likely to have masses on the order of 50 to 500 kg with up to 0.5 kW of electrical power available. A power processing unit for a 0.2 kW-class ion thruster is currently under development for these applications. The first step in this effort is the development and testing of a 0.24 kW beam power supply. The design incorporates a 20 kHz full bridge topology with multiple secondaries connected in series to obtain outputs of up to 1200 V(sub DC). A current-mode control pulse width modulation circuit built using discrete components was selected for this application. An input voltage of 28 +/- 4 V(sub DC) was assumed, since the small spacecraft for which this system is targeted are anticipated to have unregulated low voltage busses. Efficiencies in excess of 91 percent were obtained at maximum output power. The total mass of the breadboard was less than 1.0 kg and the component mass was 0.53 kg. It is anticipated that a complete flight power processor could weigh about 2.0 kg.

Pinero, Luis R.↗
NASA NTRS1991

Performance and optimization of a 'derated' ion thruster for auxiliary propulsion

This paper discusses the characteristics and implications of use of a derated ion thruster for north-south-stationkeeping (NSSK) propulsion. A derated thruster is a 30-cm diameter primary propulsion ion thruster operated at highly throttled conditions appropriate to NSSK functions. The performance characteristics of a 30-cm ion thruster are presented, emphasizing throttled operation at low specific impulse and high thrust-to-power ratio. Performance data and component erosion are compared to other NSSK ion thrusters. Operations benefits derived from the performance advantages of the derated approach are examined assuming an Intelsat VII-type spacecraft. Minimum ground test facility pumping capabilities required to maintain facility enhanced accelerator grid erosion at acceptable levels in a lifetest are quantified as a function of thruster operating condition. Novel approaches to reducing the derated thruster mass and volume are also discussed.

Patterson, Michael J.↗
NASA NTRS2007

Thermal Cycle Lifetest of Swaged Cathode Heaters

This viewgraph presentation reviews the thermal cycling test for the Dawn mission. The flight system, the mission requirements, and the Ion Propulsion System (IPS) are shown. The Dawn mission requires periodic thruster shutdown for data transmission and coast periods. The thermal cycling test is designed to simulate approximately three complete mission profiles. The results of the tests are reviewed.

ion engines↗
NASA NTRS2008

Performance Evaluation of the Prototype Model NEXT Ion Thruster

The performance testing results of the first prototype model NEXT ion engine, PM1, are presented. The NEXT program has developed the next generation ion propulsion system to enhance and enable Discovery, New Frontiers, and Flagship-type NASA missions. The PM1 thruster exhibits operational behavior consistent with its predecessors, the engineering model thrusters, with substantial mass savings, enhanced thermal margins, and design improvements for environmental testing compliance. The dry mass of PM1 is 12.7 kg. Modifications made in the thruster design have resulted in improved performance and operating margins, as anticipated. PM1 beginning-of-life performance satisfies all of the electric propulsion thruster mission-derived technical requirements. It demonstrates a wide range of throttleability by processing input power levels from 0.5 to 6.9 kW. At 6.9 kW, the PM1 thruster demonstrates specific impulse of 4190 s, 237 mN of thrust, and a thrust efficiency of 0.71. The flat beam profile, flatness parameters vary from 0.66 at low-power to 0.88 at full-power, and advanced ion optics reduce localized accelerator grid erosion and increases margins for electron backstreaming, impingement-limited voltage, and screen grid ion transparency. The thruster throughput capability is predicted to exceed 750 kg of xenon, an equivalent of 36,500 hr of continuous operation at the full-power operating condition.

Herman, Daniel A.↗
NASA NTRS2005

Development of Spacecraft to Exploit Electric Propulsion for Outer Planetary Exploration

Outer planetary exploration spacecraft have relied on Radioisotope Power Systems (RPS) to provide power. This is necessary as solar power is not useful beyond the inner planets. For propulsion these spacecraft have made use of chemical systems. A study was undertaken to look at the possibility of designing a spacecraft for outer planetary exploration that would use an RPS in combination with Electric Propulsion. That is, the RPS would be the sole power provider to the EP system. Recent improvements in RPS s have made Radioisotope Electric Propulsion (REP) more of a possibility. The combined power and propulsion technologies of REP and a direct trajectory would potentially enable a new class of missions - high delta V, beyond Mars orbit for a small spacecraft primarily suited for small body capture (orbit capture or co-orbit). A study was undertaken to evaluate the feasibility of a REP spacecraft for a selected representative mission of this type. To evaluate the potential of these technologies, a design reference mission was established and a conceptual design of a spacecraft developed. The following paper describes the Ion Propulsion System, one particular subsystem onboard of the spacecraft.

Roche, Joseph M.↗
NASA NTRS2014

Planetary Science Technology Infusion Study: Findings and Recommendations Status

The Planetary Science Division (PSD) within the National Aeronautics and Space Administrations (NASA) Science Mission Directorate (SMD) at NASA Headquarters sought to understand how to better realize a scientific return on spacecraft system technology investments currently being funded. In order to achieve this objective, a team at NASA Glenn Research Center was tasked with surveying the science and mission communities to collect their insight on technology infusion and additionally sought inputs from industry, universities, and other organizations involved with proposing for future PSD missions. This survey was undertaken by issuing a Request for Information (RFI) activity that requested input from the proposing community on present technology infusion efforts. The Technology Infusion Study was initiated in March 2013 with the release of the RFI request. The evaluation team compiled and assessed this input in order to provide PSD with recommendations on how to effectively infuse new spacecraft systems technologies that it develops into future competed missions enabling increased scientific discoveries, lower mission cost, or both. This team is comprised of personnel from the Radioisotope Power Systems (RPS) Program and the In-Space Propulsion Technology (ISPT) Program staff.The RFI survey covered two aspects of technology infusion: 1) General Insight, including: their assessment of barriers to technology infusion as related to infusion approach; technology readiness; information and documentation products; communication; integration considerations; interaction with technology development areas; cost-capped mission areas; risk considerations; system level impacts and implementation; and mission pull. 2) Specific technologies from the most recent PSD Announcements of Opportunities (AOs): The Advanced Stirling Radioisotope Generator (ASRG), aerocapture and aeroshell hardware technologies, the NASA Evolutionary Xenon Thruster (NEXT) ion propulsion system, and the Advanced Materials Bi-propellant Rocket (AMBR) engine.This report will present the teams Findings from the RFI inputs and the recommendations that arose from these findings. Methodologies on the findings and recommendations development are discussed.

technology infusion↗
NASA NTRS2017

Overview of the Development of the Advanced Electric Propulsion System (AEPS)

NASA is committed to the demonstration and application of high-power solar electric propulsion to meet its future mission needs. It is continuing to develop the 14 kW Advanced Electric Propulsion System (AEPS) under a project that recently completed an Early Integrated System Test (EIST) and System Preliminary Design Review (PDR). In addition, NASA is pursuing external partnerships in order to demonstrate Solar Electric Propulsion (SEP) technology and the advantages of high-power electric propulsion-based spacecraft. The recent announcement of a Power and Propulsion Element (PPE) as the first major piece of an evolvable human architecture to Mars has replaced the Asteroid Redirect Robotic Mission (ARRM) as the most likely first application of the AEPS Hall thruster system. This high-power SEP capability, or an extensible derivative of it, has been recognized as a critical part of a new, affordable human exploration architecture for missions beyond-low-Earth-orbit. This paper presents the status of AEPS development activities, and describes how AEPS hardware will be integrated into the PPE ion propulsion system.

Herman, Daniel↗
NASA NTRS1987

TAU - A mission to a thousand astronomical units

The potential of nuclear electric propulsion (NEP) is investigated as the enabling technology for achieving a mission to a thousand astronomical units within 50 years duration. By means of a 1000 AU baseline, the primary objective is to make measurements of distances to the stars in the Galaxy and beyond. In addition, several deep space unique studies in astronomy, astrophysics, cosmology, and space plasma physics can be carried out. NEP technology requirements for a mission to 1000 AU are addressed. These technology requirements are compared with current plans for both nuclear space power and ion propulsion research. And finally an example TAU spacecraft system is described.

Nock, K. T.↗
NASA NTRS2004

Development of In-Situ Erosion Measurement Techniques for Application to Real-Time Determination of Plasma Thruster Component Lifetimes

This research has resulted in advancing the laser-based diagnostic capability and the ion optics development in the ion propulsion program at NASA GRC. Laser-based plasma diagnostics have been demonstrated in tabletop experiments and, in the case of LDI, on laboratory hollow cathodes. Assessment by GRC of its resources and priorities two years into the grant lead to a refocusing of the research effort away from the development of a real-time erosion rate measurement technique. The extension of the diagnostic techniques to diagnostic tools has been transferred to graduate students under the technical direction of the PI. These diagnostics may facilitate the development of ion thruster with significantly improved throughput capability for lower-power (10 kW) missions High-Isp, Long-lived ion optics development has proceeded from simple extensions of state-of-the-art geometries to radically different geometries and materials. Full-scale testing of these ion optics has demonstrated a significant advance in the throughput capability of ion thrusters enabling significantly more demanding missions. The capability to predict the throughput was developed and will continue to be upgraded. The performance models have been validated via full-scale testing. Partial validation of the throughput prediction will be completed via an upcoming wear test of the ion optics.

Source record↗
NASA NTRS2018

Overview of the Development and Mission Application of the Advanced Electric Propulsion System (AEPS)

NASA remains committed to the development and demonstration of a high-power solar electric propulsion capability for the Agency. NASA is continuing to develop the 14 kW Advanced Electric Propulsion System (AEPS), which has recently completed an Early Integrated System Test and System Preliminary Design Review. NASA continues to pursue Solar Electric Propulsion (SEP) Technology Demonstration Mission partners and mature high-power SEP mission concepts. The recent announcement of the development of a Power and Propulsion Element (PPE) as the first element of an evolvable human architecture to Mars has replaced the Asteroid Redirect Robotic Mission (ARRM) as the most probable first application of the AEPS Hall thruster system. This high-power SEP capability, or an extensible derivative of it, has been identified as a critical part of an affordable, beyond-low-Earth-orbit, manned exploration architecture. This paper presents the status of the combined NASA and Aerojet Rocketdyne AEPS development activities and updated mission concept for implementation of the AEPS hardware as part of the ion propulsion system for a PPE.

Herman, Daniel A.↗
NASA NTRS1993

TROPIX: A solar electric propulsion flight experiment

The Transfer Orbit Plasma Interaction Experiment (TROPIX) is a proposed scientific experiment and flight demonstration of a solar electric propulsion vehicle. Its mission goals are to significantly increase our knowledge of Earth's magnetosphere and its associated plasma environment and to demonstrate an operational solar electric upper stage (SEUS) for small launch vehicles. The scientific investigations and flight demonstration technology experiments are uniquely interrelated because of the spacecraft's interaction with the surrounding environment. The data obtained will complement previous studies of the Earth's magnetosphere and space plasma environment by supplying the knowledge necessary to attain the strategic objectives of the NASA Office of Space Science. This first operational use of a primary ion propulsion vehicle, designed to withstand the harsh environments from low Earth orbit to geosynchronous Earth orbit, may lead to the development of a new class of electric propulsion upper stages or space-based transfer vehicles and may improve future spacecraft design and safety.

Hickman, J. Mark↗
NASA NTRS1975

A 20000-hour endurance test of a structurally and thermally integrated 5-cm diameter ion thruster main cathode

A 5-cm diameter mercury ion thruster main cathode has completed over 20,000 hours of operation in an ongoing lifetime endurance test. The cathode operating parameters remained at acceptable performance levels throughout the test, the first 9175 hours of which were part of a thruster endurance test. After 20,000 hours, the cathode discharge was easily restarted, the tip orifice indicated negligible erosion and the tip heater showed no degradation. The cathode-isolator- vaporizer assembly, a major thruster subsystem, has thus successfully demonstrated an operational lifetime capability of 20,000 hours, which is the lifetime goal of the 8-cm diameter auxiliary propulsion ion thruster.

Wintucky, E. G.↗
NASA NTRS1975

A 20,000-hour endurance test of a structurally and thermally integrated 5-cm diameter ion thruster main cathode

A 5-cm diameter mercury ion thruster main cathode has completed over 20,000 hours of operation in an ongoing lifetime endurance test. The cathode operating parameters remained at acceptable performance levels throughout the test, the first 9175 hours of which were part of a thruster endurance test. After 20,000 hours, the cathode discharge was easily restarted, the tip orifice indicated negligible erosion and the tip heater showed no degradation. The cathode-isolator-vaporizer assembly, a major thruster subsystem, has thus successfully demonstrated an operational lifetime capability of 20,000 hours, which is the lifetime goal of the 8-cm diameter auxiliary propulsion ion thruster.

Wintucky, E. G.↗
NASA NTRS2000

The Use of Environmental Test Facilities for Purposes Beyond Their Original Design

Increasing demands from space flight project offices are requiring environmental testing facilities to become more versatile with increased capabilities. At the same time, maintaining a cost-effective approach to test operations has driven efforts to use these facilities for purposes beyond their original design. This paper presents an overview of the Jet Propulsion Laboratory's efforts to provide JPL's space flight projects with test facilities to meet unique test requirements and to serve the needs of selected outside customers. The large number of recent Mars Missions, including the Mars Pathfinder project, have required testing of components and systems in a Martian surface environment in facilities originally designed for deep space testing. The unique problems associated with performing these tests are discussed, along with practical solutions. Other unique test requirements are discussed including the use of space simulation chambers for testing high altitude balloon gondolas and the use of vacuum chambers for system level test firing of an ion propulsion engine.

Fisher, Terry C.↗
NASA NTRS2009

NEXT Long-Duration Test Neutralizer Performance and Erosion Characteristics

The NASA's Evolutionary Xenon Thruster (NEXT) program is developing the next-generation ion propulsion system with significant enhancements beyond the state-of-the-art to provide future NASA science missions with enhanced capabilities at a low total development cost. A Long-Duration Test (LDT) was initiated in June 2005, to verify the NEXT propellant throughput capability to a qualification-level of 450 kg, 1.5 times the anticipated throughput requirement of 300 kg per thruster based on mission analyses. As of September 2, 2009, the thruster has accumulated 24,400 hr of operation with extensive durations at the following input powers: 6.9, 4.7, 1.1, and 0.5 kW. The thruster has processed 434 kg of xenon, surpassing the NASA Solar Technology Application Readiness (NSTAR) program thruster propellant throughput demonstrated during the extended life testing of the Deep Space 1 flight spare ion thruster and approaching the NEXT development qualification throughput goal of 450 kg. The NEXT LDT has demonstrated a total impulse of 16.1 10(exp 6zzz0 N s; the highest total impulse ever demonstrated by an ion thruster. A reduction in neutralizer flow margin has been the only appreciable source of thruster performance degradation. The behavior of the neutralizer is not easily predicted due to both erosion and deposition observed in previous wear tests. Spot-to-plume mode transition flow data and in-situ erosion results for the LDT neutralizer are discussed. This loss of flow margin has been addressed through a combination of a design change in the prototype-model neutralizer to increase flow margin at low emission current and to update the NEXT throttle table to ensure adequate flow margin as a function of propellant throughput processed. The new throttle table will be used for future LDT operations. The performance of the NEXT LDT neutralizer is consistent with that observed for long-life hollow cathodes. The neutralizer life-limiting failure modes are progressing as expected and the neutralizer data indicate none of the neutralizer failures are imminent.

Herman, Daniel A.↗
NASA NTRS2004

Space Electronic Test Engineering

The Space Power and Propulsion Test Engineering Branch at NASA Glenn Research center has the important duty of controlling electronic test engineering services. These services include test planning and early assessment of Space projects, management and/or technical support required to safely and effectively prepare the article and facility for testing, operation of test facilities, and validation/delivery of data to customer. The Space Electronic Test Engineering Branch is assigned electronic test engineering responsibility for the GRC Space Simulation, Microgravity, Cryogenic, and Combustion Test Facilities. While working with the Space Power and Propulsion Test Engineering Branch I am working on several different assignments. My primary assignment deals with an electrical hardware unit known as Sunny Boy. Sunny Boy is a DC load Bank that is designed for solar arrays in which it is used to convert DC power form the solar arrays into AC power at 60 hertz to pump back into the electricity grid. However, there are some researchers who decided that they would like to use the Sunny Boy unit in a space simulation as a DC load bank for a space shuttle or even the International Space Station hardware. In order to do so I must create a communication link between a computer and the Sunny Boy unit so that I can preset a few of the limits (such power, set & constant voltage levels) that Sunny Boy will need to operate using the applied DC load. Apart from this assignment I am also working on a hi-tech circuit that I need to have built at a researcher s request. This is a high voltage analog to digital circuit that will be used to record data from space ion propulsion rocket booster tests. The problem that makes building this circuit so difficult is that it contains high voltage we must find a way to lower the voltage signal before the data is transferred into the computer to be read. The solution to this problem was to transport the signal using infrared light which will lower the voltage signal down low enough so that it is harmless to a computer. Along with my involvement in the Space Power and Propulsion Test Engineering Branch, I am obligated to assist all other members of the branch in their work. This will help me to strengthen and extend my knowledge of Electrical Engineering.

Chambers, Rodney D.↗