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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 559 records · Page 31
All sourcesNASA NTRSDOE OSTI
NASA NTRS2003

An overview of the Nuclear Electric Xenon Ion System (NEXIS) program

NASA is investigating high power, high specific impulse propulsion technologies that could enable ambitious flights such as multi-body rendezvous missions, outer planet orbiters and interstellar precursor missions. The requirements for these missions are much more demanding than those for state-of-the-art solar-powered ion propulsion applications. The purpose of the NEXIS program is to develop advanced ion thruster technologies that satisfy the requirements for high power, high specific impulse operation, high efficiency and long thruster life. The nominal design point for the NEXIS thruster is 20 kWe at a specific impulse of 7500 s with an efficiency over 78% and a xenon throughput capability of greater than 2000 kg. These performance and throughput goals will be achieved by applying a combination of advanced technologies including a large discharge chamber, erosion resistant carbon-carbon grids, an advanced reservoir hollow cathode and techniques for increasing propellant efficiency such as grid masking and accelerator grid aperture diameter tailoring. This paper provides an overview of the challenges associated with these requirements and how they are being addressed in the NEXIS program.

carbon-carbon grids↗
NASA NTRS1990

Xenon ion sources for space applications

The technology of xenon ion thrusters under development for space-propulsion and plasma-contactor applications is described with particular consideration given to the 1.4-kW Xenon Ion Propulsion Subsystem (XIPS) designed for the upcoming Atmospheric Laboratory for Applications and Science Shuttle-based missions. The 25-cm-diam ion thruster of the XIPS is able to convert input power into thrust-beam power with an efficiency of over 82 percent. The XIPS will be used for satellite stationkeeping and orbit raising and maneuvering and as a plasma contactor. A 1991 flight application is described in which the 25-cm XIPS plasma contactor will be used to stabilize the electrical potential of the Shuttle Orbiter during firings of a 1.5-A electron beam from its payload bay.

Beattie, J. R.↗
NASA NTRS2017

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.

xenon↗
NASA NTRS1989

Electric propulsion options for 10 kW class earth space missions

Five and 10 kW ion and arcjet propulsion system options for a near-term space demonstration experiment have been evaluated. Analyses were conducted to determine first-order propulsion system performance and system component mass estimates. Overall mission performance of the electric propulsion systems was quantified in terms of the maximum thrusting time, total impulse, and velocity increment capability available when integrated onto a generic spacecraft under fixed mission model assumptions. Maximum available thrusting times for the ion-propelled spacecraft options, launched on a DELTA II 6920 vehicle, range from approximately 8,600 hours for a 4-engine 10 kW system to more than 29,600 hours for a single-engine 5 kW system. Maximum total impulse values and maximum delta-v's range from 1.2x10(7) to 2.1x10(7) N-s, and 3550 to 6200 m/s, respectively. Maximum available thrusting times for the arcjet propelled spacecraft launched on the DELTA II 6920 vehicle range from approximately 528 hours for the 6-engine 10 kW hydrazine system to 2328 hours for the single-engine 5 kW system. Maximum total impulse values and maximum delta-v's range from 2.2x10(6) to 3.6x10(6) N-s, and approximately 662 to 1072 m/s, respectively.

Patterson, M. J.↗
NASA NTRS1977

Space processing applications of ion beam technology

Ion thruster engines for spacecraft propulsion can serve as ion beam sources for potential space processing applications. The advantages of space vacuum environments and the possible gravity effects on thruster ion beam materials operations such as thin film growth, ion milling, and surface texturing were investigated. The direct gravity effect on sputter deposition and vapor deposition processes are discussed as well as techniques for cold and warm welding.

Grodzka, P. G.↗
NASA NTRS1986

Multimegawatt nuclear electric propulsion - First order system design and performance evaluation

The development of space nuclear power supplies capable of delivering high levels of power (1 to 10 MWe) for long life (5 to 10 years) is being considered. An important potential application of these systems will be to provide power for advanced electric propulsion devices such as ion thrusters or Magneto-Plasma-Dynamic (MPD) thrusters. High-powered Nuclear Electric Propulsion (NEP) concepts based on both ion and MPD thrusters are reviewed. To allow for mission performance evaluation, power supply, power processing, and thruster subsystems are characterized for both ion and MPD propulsion systems designed for operation in the 1 to 10 MWe power range. Several proposed missions which are significantly enhanced by the performance potential of these devices are investigated. Studied missions include near-earth orbit raising, cis-lunar cargo transportation, and planetary exploration. If developed, these 1 to 10 MWe NEP systems may provide a technological base for the development of even larger propulsion systems with power levels of up to hundreds of megawatts. The potential performance and technologies associated with such ultra-high-powered NEP systems is explored.

Sercel, J. C.↗
NASA NTRS1989

Electric Propulsion Options for 10 kW Class Earth-Space Missions

Five and 10 kW ion and arcjet propulsion system options for a near-term space demonstration experiment were evaluated. Analyses were conducted to determine first-order propulsion system performance and system component mass estimates. Overall mission performance of the electric propulsion systems was quantified in terms of the maximum thrusting time, total impulse, and velocity increment capability available when integrated onto a generic spacecraft under fixed mission model assumptions. Maximum available thrusting times for the ion-propelled spacecraft options, launched on a DELTA 2 6920 vehicle, range from approximately 8,600 hours for a 4-engine 10 kW system to more than 29,600 hours for a single-engine 5 kW system. Maximum total impulse values and maximum delta-v's range from 1.2x10 (exp 7) to 2.1x10 (exp 7) N-s, and 3550 to 6200 m/s, respectively. Maximum available thrusting times for the arcjet propelled spacecraft launched on the DELTA 2 6920 vehicle range from approximately 528 hours for the 6-engine 10 kW hydrazine system to 2328 hours for the single-engine 5 kW system. Maximum total impulse values and maximum delta-v's range from 2.2x10 (exp 6) to 3.6x10 (exp 6) N-s, and approximately 662 to 1072 m/s, respectively.

Patterson, M. J.↗
DOE OSTI2024

CubeSat Interplanetary Exploration: A Deep Dive into Nuclear Propulsion and Astrodynamics for Small Satellites

CubeSats are small satellites most commonly used by universities, however, nowadays places like NASA are looking to CubeSats for interplanetary missions. Because of their universal form fit it makes them cost effective, economical, and therefor desirable for deep space. To achieve deep space exploration, we must explore the applications of propulsion and astrodynamics on small satellites. This paper will touch on one scaling Ion and chemical propulsion as well as nuclear kilo reactors. Two choreographing a flight path for these small satellites and lastly designing an efficient cost-effective platform for CubeSats to get to deep space territory.

22 GENERAL STUDIES OF NUCLEAR REACTORS↗
NASA NTRS1986

Advanced propulsion for polar-orbiting and coorbiting free flyers

Advanced propulsion systems for polar-orbiting and coorbiting free flyers were investigated. Resistojet, arcjet, ion, magnetoplasmadynamic and chemical-bipropellant nitrogen tetroxide/monomethyl hydrazine (NTO/MMH) propulsion systems were compared to the baseline free-flyer hydrazine (N2H4)-propulsion-system performance. Advanced resistojet, arcjet, ion, and NTO/MMH propulsion systems enable significant propellant-mass savings over the baseline N2H4-propulsion system. Using free-flyer mission requirements from the Langley Research Center Mission-Data Base, detailed propulsion requirements for over thirty free-flyer missions were analyzed. The Polar-Platform trip-time constraints may preclude using a low-thrust electric-propulsion system. Electric propulsion will, however, allow a significant coorbiter propellant-mass reduction. Frequent servicing and nodal-regression effects on the coorbiting free-flyer's orbit increase the required mission velocity change and propellant mass. For many coorbiter missions high-specific-impulse resistojet-, arcjet- and ion-propulsion systems allow substantial life-cycle propellant-mass savings.

Palaszewski, B. A.↗
NASA NTRS2004

Ultra High Voltage Propellant Isolators and Insulators for JIMO Ion Thrusters

Within NASA's Project Prometheus, high specific impulse ion thrusters for electric propulsion of spacecraft for the proposed Jupiter Icy Moon Orbiter (JIMO) mission to three of Jupiter's moons: Callisto, Ganymede and Europa will require high voltage operation to meet mission propulsion. The anticipated approx.6,500 volt net ion energy will require electrical insulation and propellant isolation which must exceed that used successfully by the NASA Solar Electric Propulsion Technology Readiness (NSTAR) Deep Space 1 mission thruster by a factor of approx.6. Xenon propellant isolator prototypes that operate at near one atmosphere and prototypes that operate at low pressures (<100 Torr) have been designed and are being tested for suitability to the JIMO mission requirements. Propellant isolators must be durable to Paschen breakdown, sputter contamination, high temperature, and high voltage while operating for factors longer duration than for the Deep Space 1 Mission. Insulators used to mount the thrusters as well as those needed to support the ion optics have also been designed and are under evaluation. Isolator and insulator concepts, design issues, design guidelines, fabrication considerations and performance issues are presented. The objective of the investigation was to identify candidate isolators and insulators that are sufficiently robust to perform durably and reliably during the proposed JIMO mission.

Banks, Bruce A.↗
NASA NTRSDate not supplied

Design Study of a Nuclear-Electric Transport Vehicle for Ice Giant Missions

Nuclear-electric propulsion, in which a nuclear reactor is used to power high specific-impulse ion thrusters for propulsion, is a promising method for such applications as missions to the ice giants Uranus and Neptune, identified as targets of high science value in the most recent NASA Decadal Survey of Planetary Science. Until recently, however, this has required a lightweight space nuclear power system that has yet to be designed. NASA has recently been developing the “Kilopower” nuclear reactor as a power source for future exploration missions, with a 1-kW prototype recently tested under the KRUSTY program. We utilized a next-generation Kilopower reactor to design a transportation system for missions to the ice giant planets. The mission objectives were to carry a payload of 1164 kg, comprising a lander/hopper vehicle to be delivered to explore the moon or moons of the target ice giant, along with atmospheric probes to drop into the atmosphere, plus 188 kg of science instrumentation on the spacecraft itself.

nuclear electric propulsion↗
NASA NTRS1975

A structural and thermal packaging approach for power processing units for 30-cm ion thrusters

Solar Electric Propulsion (SEP) is currently being studied for possible use in a number of near earth and planetary missions. The thruster subsystem for these missions would consist of 30 centimeter ion thrusters with Power Processor Units (PPU) clustered in assemblies of from two to ten units. A preliminary design study of the electronic packaging of the PPU has been completed at Lewis Research Center of NASA. This study evaluates designs meeting the competing requirements of low system weight and overall mission flexibility. These requirements are evaluated regarding structural and thermal design, electrical efficiency, and integration of the electrical circuits into a functional PPU layout.

Maloy, J. E.↗
NASA NTRS1975

A structural and thermal packaging approach for power processing units for 30-cm ion thrusters

Solar Electric Propulsion (SEP) is currently being studied for possible use in a number of near earth and planetary missions. The thruster subsystem for these missions would consist of 30 centimeter ion thrusters with Power Processor Units (PPU) clustered in assemblies of from two to ten units. A preliminary design study of the electronic packaging of the PPU has been completed at Lewis Research Center of NASA. This study evaluates designs meeting the competing requirements of low system weight and overall mission flexibility. These requirements are evaluated regarding structural and thermal design, electrical efficiency, and integration of the electrical circuits into a functional PPU layout.

Maloy, J. E.↗
NASA NTRS2004

Experimental Investigations from the Operation of a 2 Kw Brayton Power Conversion Unit and a Xenon Ion Thruster

A 2 kW Brayton Power Conversion Unit (PCU) and a xenon ion thruster were integrated with a Power Management and Distribution (PMAD) system as part of a Nuclear Electric Propulsion (NEP) Testbed at NASA's Glenn Research Center. Brayton converters and ion thrusters are potential candidates for use on future high power NEP missions such as the proposed Jupiter Icy Moons Orbiter (JIMO). The use of existing lower power test hardware provided a cost-effective means to investigate the critical electrical interface between the power conversion system and ion propulsion system. The testing successfully demonstrated compatible electrical operations between the converter and the thruster, including end-to-end electric power throughput, high efficiency AC to DC conversion, and thruster recycle fault protection. The details of this demonstration are reported herein.

Mason, Lee↗
NASA NTRS1977

Ion Drive multi-asteroid rendezvous trajectories

The general characteristics of multiasteroid rendezvous missions are examined, taking into account a use of an improved solar electric propulsion system, called Ion Drive. The improvements are related to the development of light-weight solar cells, extensions to high thruster power levels with high exhaust velocity, and the attachment of solar concentrators to increase the power available beyond 1.5 AU. Flight times on the order of 500-800 days are generally required. It is assumed that a stay time of about 60 days at each asteroid is adequate. Attention is given to launch and spacecraft parameters, the search technique, and multiasteroid tours involving Ceres and Vesta. It is found that at least four asteroids in the main belt, generally selected from those with known characteristics, may be reached on a single mission with a 100-kg separable probe landed on each one.

Bender, D. F.↗
NASA NTRS1987

Nuclear reactor power for an electrically powered orbital transfer vehicle

To help determine the systems requirements for a 300-kWe space nuclear reactor power system, a mission and spacecraft have been examined which utilize electric propulsion and this nuclear reactor power for multiple transfers of cargo between low earth orbit (LEO) and geosynchronous earth orbit (GEO). A propulsion system employing ion thrusters and xenon propellant was selected. Propellant and thrusters are replaced after each sortie to GEO. The mass of the Orbital Transfer Vehicle (OTV), empty and dry, is 11,000 kg; nominal propellant load is 5000 kg. The OTV operates between a circular orbit at 925 km altitude, 28.5 deg inclination, and GEO. Cargo is brought to the OTV by Shuttle and an Orbital Maneuvering Vehicle (OMV); the OTV then takes it to GEO. The OTV can also bring cargo back from GEO, for transfer by OMV to the Shuttle. OTV propellant is resupplied and the ion thrusters are replaced by the OMV before each trip to GEO. At the end of mission life, the OTV's electric propulsion is used to place it in a heliocentric orbit so that the reactor will not return to earth. The nominal cargo capability to GEO is 6000 kg with a transit time of 120 days; 1350 kg can be transferred in 90 days, and 14,300 kg in 240 days. These capabilities can be considerably increased by using separate Shuttle launches to bring up propellant and cargo, or by changing to mercury propellant.

Jaffe, L.↗
NASA NTRS1988

Nuclear-electric reusable orbital transfer vehicle

To help determine the systems requirements for a 300-kWe space nuclear reactor power system, a mission and spacecraft have been examined that utilize electric propulsion supported by the nuclear reactor's power for multiple transfers of cargo between low earth orbit (LEO) and geosynchronous earth orbit (GEO). A propulsion system employing ion thrusters and xenon propellant was selected. Propellant and thrusters are replaced after each sortie to GEO. The mass of the orbital transfer vehicle (OTV), empty and dry, is 11,000 kg; nominal propellant load is 5000 kg. The OTV operates between a circular orbit at 925-km altitude, 28.5-deg inclination, and GEO. Cargo is brought to the OTV by Shuttle and an orbital maneuvering vehicle (OMV); the OTV then takes it to GEO. The OTV can also bring cargo back from GEO for transfer by OMV to the Shuttle. OTV propellant is resupplied, and the ion thrusters are replaced, by the OMV before each sortie to GEO. At the end of mission life, the OTV's electric propulsion is used to place it in a heliocentric orbit so that the reactor will not return to earth. The nominal cargo capability to GEO is 6000 kg, with a transit time of 120 days; 1350 kg can be transferred in 90 days, and 14,300 kg in 240 days. These capabilities can be considerably increased by using separate Shuttle launches to bring up propellant and cargo or by changing to mercury propellant.

Jaffe, Leonard D.↗