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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 NTRS2004

Performance and Vibration of 30 cm Pyrolytic Ion Thruster Optics

Carbon has a sputter erosion rate about an order of magnitude less than that of molybdenum, over the voltages typically used in ion thruster applications. To explore its design potential, 30 cm pyrolytic carbon ion thruster optics have been fabricated geometrically similar to the molybdenum ion optics used on NSTAR. They were then installed on an NSTAR Engineering Model thruster, and experimentally evaluated over much of the original operating envelope. Ion beam currents ranged from 0.51 to 1.76 Angstroms, at total voltages up to 1280 V. The perveance, electron back-streaming limit, and screen-grid transparency were plotted for these operating points, and compared with previous data obtained with molybdenum. While thruster performance with pyrolytic carbon was quite similar to that with molybdenum, behavior variations can reasonably be explained by slight geometric differences. Following all performance measurements, the pyrolytic carbon ion optics assembly was subjected to an abbreviated vibration test. The thruster endured 9.2 g(sub rms) of random vibration along the thrust axis, similar to DS 1 acceptance levels. Despite significant grid clashing, there was no observable damage to the ion optics assembly.

Haag, Thomas↗
NASA NTRS2015

20-mN Variable Specific Impulse (Isp) Colloid Thruster

Busek Company, Inc., has designed and manufactured an electrospray emitter capable of generating 20 mN in a compact package (7x7x1.7 in). The thruster consists of nine porous-surface emitters operating in parallel from a common propellant supply. Each emitter is capable of supporting over 70,000 electrospray emission sites with the plume from each emitter being accelerated through a single aperture, eliminating the need for individual emission site alignment to an extraction grid. The total number of emission sites during operation is expected to approach 700,000. This Phase II project optimized and characterized the thruster fabricated during the Phase I effort. Additional porous emitters also were fabricated for full-scale testing. Propellant is supplied to the thruster via existing feed-system and microvalve technology previously developed by Busek, under the NASA Space Technology 7's Disturbance Reduction System (ST7-DRS) mission and via follow-on electric propulsion programs. This project investigated methods for extending thruster life beyond the previously demonstrated 450 hours. The life-extending capabilities will be demonstrated on a subscale version of the thruster.

Demmons, Nathaniel↗
NASA NTRS2018

Development of High Power Hall Thruster Systems to Enable the NASA Exploration Vision

The next phase of space exploration missions requires high power Solar Electric Propulsion (SEP) systems for large-scale science missions and cargo transportation. Development is underway at Aerojet Rocketdyne on Hall thruster systems that are intended to bracket the needs of future NASA SEP missions in support of space exploration. The Advanced Electric Propulsion System (AEPS) program is developing and qualifying a 13.3kW Hall thruster system to be demonstrated on the Power and Propulsion Element (PPE), which is intended to be the first element of a Lunar Outpost Platform - Gateway (LOP-G). The NextSTEP program is integrating a nested Hall thruster into a 100kW system and testing it for 100 hours. These two programs will provide a path to efficient in-space propulsion that will allow NASA to transfer the large amounts of cargo that is needed to support human missions - first to the moon and then on to Mars. The Advanced Electric Propulsion System (AEPS) program is completing development, qualification and delivery of five flight 13.3kW EP systems to NASA. The flight AEPS system includes a magnetically shielded long-life Hall thruster, Power Processing Unit (PPU) and a Xenon Flow Controller (XFC). The Hall thruster, developed and demonstrated by NASA, operates at input powers up to 12.5kW while providing a specific impulse over an estimated 2800s at an input voltage of 600V. The power processor is designed to accommodate an input voltage range of 95-140V, consistent with operation beyond the orbit of Mars. The integrated system input power is continuously throttleable between 3 and 13.3kW. Component level testing of the EP String has begun with prototype hardware. The NextSTEP program is developing a 100kW Electric Propulsion (EP) system using a nested Hall thruster designed for powers up to 250kW, a modular power processor and a modular mass flow controller. While the program objective is to operate the integrated EP system continuously at 100kW for 100hrs to demonstrate thermal stability and support the development of system life time models, it builds on decades of experience with long-life Hall thrusters and the design is evolvable to a capability of 250kW. Design upgrades that demonstrate the 100kW EP system have been completed and tested. Aerojet Rocketdyne (AR) is excited to support NASA as it extends human reach into deep space and believes that these programs will provide the propulsion to make such missions affordable and sustainable. These systems provide NASA with a range of options to power its deep space transport vehicles. This paper presents the mission requirements for supporting the NASA exploration vision, as well as the status for the high power Hall thruster systems in development.

Jackson, Jerry↗
NASA NTRS2018

Development of High Power Hall Thruster Systems to Enable the NASA Exploration Vision

The next phase of space exploration missions requires high power Solar Electric Propulsion (SEP) systems for large-scale science missions and cargo transportation. Development is underway at Aerojet Rocketdyne on Hall thruster systems that are intended to bracket the needs of future NASA SEP missions in support of space exploration. The Advanced Electric Propulsion System (AEPS) program is developing and qualifying a 13.3kW Hall thruster system to be demonstrated on the Power and Propulsion Element (PPE), which is intended to be the first element of a Lunar Outpost Platform - Gateway (LOP-G). The NextSTEP program is integrating a nested Hall thruster into a 100 kW system and testing it for 100 hours. These two programs will provide a path to efficient in-space propulsion that will allow NASA to transfer the large amounts of cargo that is needed to support human missions - first to the moon and then on to Mars. The Advanced Electric Propulsion System (AEPS) program is completing development, qualification and delivery of five flight 13.3kW EP systems to NASA. The flight AEPS system includes a magnetically shielded long-life Hall thruster, Power Processing Unit (PPU) and a Xenon Flow Controller (XFC). The Hall thruster, developed and demonstrated by NASA, operates at input powers up to 12.5 kW while providing a specific impulse over an estimated 2800s at an input voltage of 600V. The power processor is designed to accommodate an input voltage range of 95-140V, consistent with operation beyond the orbit of Mars. The integrated system input power is continuously throttleable between 3 and 13.3kW. Component level testing of the EP String has begun with prototype hardware. The NextSTEP program is developing a 100kW Electric Propulsion (EP) system using a nested Hall thruster designed for powers up to 250kW, a modular power processor and a modular mass flow controller. While the program objective is to operate the integrated EP system continuously at 100kW for 100 hours to demonstrate thermal stability and support the development of system life time models, it builds on decades of experience with long-life Hall thrusters and the design is evolvable to a capability of 250kW. Design upgrades that demonstrate the 100kW EP system have been completed and tested. Aerojet Rocketdyne is excited to support NASA as it extends human reach into deep space and believes that these programs will provide the propulsion to make such missions affordable and sustainable. These systems provide NASA with a range of options to power its deep space transport vehicles. This paper presents the mission requirements for supporting the NASA exploration vision, as well as the status for the high power Hall thruster systems in development.

Jackson, Jerry↗
NASA NTRS1975

Power processing for electric propulsion

The inclusion of electric thruster systems in spacecraft design is considered. The propulsion requirements of such spacecraft dictate a wide range of thruster power levels and operational lifetimes, which must be matched by lightweight, efficient, and reliable thruster power processing systems. Electron bombardment ion thruster requirements are presented, and the performance characteristics of present power processing systems are reviewed. Design philosophies and alternatives in areas such as inverter type, arc protection, and control methods are discussed along with future performance potentials for meeting goals in the areas of power process or weight (10 kg/kW), efficiency (approaching 92 percent), reliability (0.96 for 15,000 hr), and thermal control capability (0.3 to 5 AU).

Finke, R. C.↗
NASA NTRS2013

High-Efficiency Nested Hall Thrusters for Robotic Solar System Exploration

This work describes the scaling and design attributes of Nested Hall Thrusters (NHT) with extremely large operational envelopes, including a wide range of throttleability in power and specific impulse at high efficiency (>50%). NHTs have the potential to provide the game changing performance, powerprocessing capabilities, and cost effectiveness required to enable missions that cannot otherwise be accomplished. NHTs were first identified in the electric propulsion community as a path to 100- kW class thrusters for human missions. This study aimed to identify the performance capabilities NHTs can provide for NASA robotic and human missions, with an emphasis on 10-kW class thrusters well-suited for robotic exploration. A key outcome of this work has been the identification of NHTs as nearly constant-efficiency devices over large power throttling ratios, especially in direct-drive power systems. NHT systems sized for robotic solar system exploration are predicted to be capable of high-efficiency operation over nearly their entire power throttling range. A traditional Annular Hall Thruster (AHT) consists of a single annular discharge chamber where the propellant is ionized and accelerated. In an NHT, multiple annular channels are concentrically stacked. The channels can be operated in unison or individually depending on the available power or required performance. When throttling an AHT, performance must be sacrificed since a single channel cannot satisfy the diverse design attributes needed to maintain high thrust efficiency. NHTs can satisfy these requirements by varying which channels are operated and thereby offer significant benefits in terms of thruster performance, especially under deep power throttling conditions where the efficiency of an AHT suffers since a single channel can only operate efficiently (>50%) over a narrow power throttling ratio (3:1). Designs for 10-kW class NHTs were developed and compared with AHT systems. Power processing systems were considered using either traditional Power Processing Units (PPU) or Direct Drive Units (DDU). In a PPU-based system, power from the solar arrays is transformed from the low voltage of the arrays to the high voltage needed by the thruster. In a DDU-based system, power from the solar arrays is fed to the thruster without conversion. DDU-based systems are attractive for their simplicity since they eliminate the most complex and expensive part of the propulsion system. The results point to the strong potential of NHTs operating with either PPUs or DDUs to benefit robotic and human missions through their unprecedented power and specific impulse throttling capabilities. NHTs coupled to traditional PPUs are predicted to offer high-efficiency (>50%) power throttling ratios 320% greater than present capabilities, while NHTs with direct-drive power systems (DDU) could exceed existing capabilities by 340%. Because the NHT-DDU approach is implicitly low-cost, NHT-DDU technology has the potential to radically reduce the cost of SEP-enabled NASA missions while simultaneously enabling unprecedented performance capability.

Hofer, Richard R.↗
NASA NTRS2005

Ion Engine Plume Interaction Calculations for Prototypical Prometheus 1

Prometheus 1 is a conceptual mission to demonstrate the use of atomic energy for distant space missions. The hypothetical spacecraft design considered in this paper calls for multiple ion thrusters, each with considerably higher beam energy and beam current than have previously flown in space. The engineering challenges posed by such powerful thrusters relate not only to the thrusters themselves, but also to designing the spacecraft to avoid potentially deleterious effects of the thruster plumes. Accommodation of these thrusters requires good prediction of the highest angle portions of the main beam, as well as knowledge of clastically scattered and charge exchange ions, predictions for grid erosion and contamination of surfaces by eroded grid material, and effects of the plasma plume on radio transmissions. Nonlinear interactions of multiple thrusters are also of concern. In this paper we describe two- and three-dimensional calculations for plume structure and effects of conceptual Prometheus 1 ion engines. Many of the techniques used have been validated by application to ground test data for the NSTAR and NEXT ion engines. Predictions for plume structure and possible sputtering and contamination effects will be presented.

Mandell, Myron J.↗
NASA NTRS1980

Low thrust vehicle concept study

Low thrust chemical (hydrogen-oxygen) propulsion systems configured specifically for low acceleration orbit transfer of large space systems were defined. Results indicate that it is cost effective and least risk to combine the OTV and stowed spacecraft in a single 65 K Shuttle. The study shows that the engine for an optimized low thrust stage (1) does not require very low thrust; (2) 1-3 K thrust range appears optimum; (3) thrust transient is not a concern; (4) throttling probably not worthwhile; and (5) multiple thrusters complicate OTV/LSS design and aggravate LSS loads. Regarding the optimum vehicle for low acceleration missions, the single shuttle launch (LSS and expendable OTV) is most cost effective and least risky. Multiple shuttles increase diameter 20%. The space based radar structure short OTV (which maximizes space available for packaged LSS) favors use of torus tank. Propellant tank pressures/vapor residuals are little affected by engine thrust level or number of burns.

Source record↗
NASA NTRS1983

Experimental Investigation of a Hall-Current Accelerator

The Hall-current accelerator is being investigated for use in the 1000-2000 sec. range of specific impulse. Three models of this thruster were tested. The first two models had three permanent magnets to supply the magnetic field and the third model had six magnets to supply the field. The third model thus had approximately twice the magnetic field of the first two. The first and second models differ only in the shape of the magnetic field. All other factors remained the same for the three models except for the anode-cathode distance, which was changed to allow for the three thrusters to have the same magnetic field integral between the anode and the cathode. These Hall thrusters were tested to determine the plasma properties, the beam characteristics, and the thruster characteristics. The thruster operated in three modes: (1) main cathode only, (2) main cathode with neutralizer cathode, and (3) neutralizer cathode only. The plasma properties were measured along an axial line, 1 mm inside the cathode radius, at a distance of 0.2 to 6.2 cm from the anode. Results show that the current used to heat the cathode produced nonuniformities in the magnetic field, hence also in the plasma properties. In a Hall thruster this general design appears to provide the most thrust when operated at a magnetic field less than the maximum value studied.

Plank, G. M.↗
NASA NTRS1984

Advanced solar electric propulsion (ASEP) for planetary missions

This paper defines an advanced SEP concept based upon demonstrated ion thruster technology improvements and projected improvements in solar array technology. The ASEP concept utilizes an 'integrated' rather than a 'stage' approach and assumes 30 cm Hg ion thrusters capable of operation near 7 kWe each. The mission performance using the ASEP system is presented for three missions: (1) Saturn orbiter, (2) comet rendezvous and (3) comet sample return. The ASEP system provides significant mission performance improvements. The simpler, fewer thruster ASEP system design may provide cost as well as performance benefits for the missions examined when compared to ballistic or SEP stage mission options.

Jones, R. M.↗
NASA NTRS1990

Manual control aspects of Space Station docking maneuvers

Due to an increase in spacecraft traffic forecasted for the Space Station era, researchers are investigating manual control and other aspects of docking operations with hopes of increasing safety, productivity, and likelihood of success while decreasing cost. Experiments have been performed which revealed the effect of approach velocity, in-flight anomalies, and control mode. Displays have been designed to enable flight planners to more easily overcome the difficulties presented by orbital mechanics. Improved understanding of human factors in the docking mission and other orbital maneuvers will play a significant role in design tradeoffs concerning thruster size, docking fixture style and mass, and on-board trajectory planning displays. Incorporating both empirical and analytic results into current and future planning of missions occurring not only in earth orbit, but also for missions in lunar and Mars orbit, will expand the performance envelopes of the astronauts who participate in these missions.

Brody, Adam R.↗
NASA NTRS1990

A 10,000-hr life test of an engineering model resistojet

One of the major issues associated with using resistojet thrusters on Space Station Freedom is the long life required. An engineering model resistojet was life-tested to determine if it was capable of meeting that requirement. This thruster, which was designed for 10,000 hr of operation at 2552.4 F (1400 C) or less under cyclical thermal conditions, successfully operated for 10,036 hr at 1836 F (1002 C) while undergoing 141 thermal cycles.

Slutz, Rodger J.↗
NASA NTRS1991

Analysis of nozzle lip and backflow expansion of a small hydrazine thruster

Results are presented for computations made with the direct simulation Monte Carlo method for the expansion of the thick boundary layer of a small nozzle. The objective in the investigation is to assess the boundary-layer characteristics that most affect the flow properties of the expanding gas. In the computations, large degrees of species separation and thermal nonequilibrium are observed as the gas expands around the nozzle lip. These aspects of the flow indicate the necessity of treating the problem through a discrete particle approach. Assessment is made of the sensitivity of the calculated results to the form of the boundary layer initially assumed. It is found that the amount of backflow is related to the flow temperature close to the wall. Therefore, the opportunity presents itself for the reduction of the impingement potential of such thrusters through appropriate design considerations. Experimental procedures for verification of the type of computations undertaken in the current study are discussed. It is proposed that the calculations made in the backflow region offer the best opportunity.

Boyd, Iain D.↗
NASA NTRS1998

Rocketdyne Advanced Reusable Technologies Program Update

NASA Marshall Space Flight Center Advanced Reusable Technologies/Rocket Based Combined Cycle (RBCC) program has enabled: (1) the creation of unique trajectory simulation facility; (2) development of basic operability and performance data; (3) results to date: (a) rocket thrusters performing as designed (P(sub c), MR, q, etc); (b) AAR performance data measured at M = 0 and 3; (c) rocket mode performance at simulated vacuum conditions; and (d) data reduction in progress; (4) completing testing; M = 0, 3 and trajectory M = 3 to 3.7; and (5) next logical steps: build flight type engine and flight demonstrate RBCC.

Ratekin, Gary↗
NASA NTRS2005

Development of Ionic Liquid Monopropellants for In-Space Propulsion

A family of new, low toxicity, high energy monopropellants is currently being evaluated at NASA Marshall Space Flight Center for in-space rocket engine applications such as reaction control engines. These ionic liquid monopropellants, developed in recent years by the Air Force Research Laboratory, could offer system simplification, less in-flight thermal management, and reduced handling precautions, while increasing propellant energy density as compared to traditional storable in-space propellants such as hydrazine and nitrogen tetroxide. However, challenges exist in identifying ignition schemes for these ionic liquid monopropellants, which are known to burn at much hotter combustion temperatures compared to traditional monopropellants such as hydrazine. The high temperature combustion of these new monopropellants make the use of typical ignition catalyst beds prohibitive since the catalyst cannot withstand the elevated temperatures. Current research efforts are focused on monopropellant ignition and burn rate characterization, parameters that are important in the fundamental understanding of the monopropellant behavior and the eventual design of a thruster. Laboratory studies will be conducted using alternative ignition techniques such as laser-induced spark ignition and hot wire ignition. Ignition delay, defined as the time between the introduction of the ignition source and the first sign of light emission from a developing flame kernel, will be measured using Schlieren visualization. An optically-accessible liquid monopropellant burner will be used to determine propellant burn rate as a function of pressure and initial propellant temperature. The burn rate will be measured via high speed imaging through the chamber s windows.

Blevins, John A.↗
NASA NTRS2015

Proof-of-Concept Experiments on a Gallium-Based Ignitron for Pulsed Power Applications

Ignitrons are electrical switching devices that operate at switching times that are on the order of microseconds, can conduct high currents of thousands of amps, and are capable of holding off tens of thousands of volts between pulses. They consist of a liquid metal pool within an evacuated tube that serves both the cathode and the source of atoms and electrons for an arc discharge. Facing the liquid metal pool is an anode suspended above the cathode, with a smaller ignitor electrode tip located just above the surface of the cathode. The ignitron can be charged to significant voltages, with a potential difference of thousands of volts between anode and cathode. When an ignition pulse is delivered from the ignitor electrode to the cathode, a small amount of the liquid metal is vaporized and subsequently ionized, with the high voltage between the anode and cathode causing the gas to bridge the gap between the two electrodes. The electrons and ions move rapidly towards the anode and cathode, respectively, with the ions liberating still more atoms from the liquid metal cathode surface as a high-current plasma arc discharge is rapidly established. This arc continues in a self-sustaining fashion until the potential difference between the anode and cathode drops below some critical value. Ignitrons have been used in a variety of pulsed power applications, including the railroad industry, industrial chemical processing, and high-power arc welding. In addition, they might prove useful in terrestrial power grid applications, serving as high-current fault switches, quickly shunting dangerous high-current or high-voltage spikes safely to ground. The motivation for this work stemmed from the fact that high-power, high-reliability, pulsed power devices like the ignitron have been used for ground testing in-space pulsed electric thruster technologies, and the continued use of ignitrons could prove advantageous to the future development and testing of such thrusters. Previous ignitron designs have used mercury as the liquid metal cathode, owing to its presence as a liquid at room temperatures and a vapor pressure of 10 Pa (75 mtorr) at room temperature. While these are favorable properties, there are obvious environmental and personal safety concerns with the storage, handling, and use of mercury and its compounds. The purpose of the present work was to fabricate and test an ignitron that used as its cathode an alternate liquid metal that was safe to handle and store. To that end, an ignitron test article that used liquid gallium as the cathode material was developed and tested. Gallium is a metal that has a melting temperature of 29.76 C, which is slightly above room temperature, and a boiling point of over 2,300 C at atmospheric pressure. This property makes gallium the element with the largest relative difference between melting and boiling points. Gallium has a limited role in biology, and when ingested, it will be subsequently processed by the body and expelled rather than accumulating to toxic levels. The next section of this Technical Memorandum (TM) provides background information on the development of mercury-based ignitrons, which serves as the starting point for the development of the gallium-based variant. Afterwards, the experimental hardware and setup used in proof-of-concept testing of a basic gallium ignitron are presented. Experimental data, consisting of discharge voltage and current waveforms as well as high-speed imaging of the gallium arc discharge in the gallium ignitron test article, are presented to demonstrate the efficacy of the concept. Discussion of the data and suggestions on improvements for future iterations of the design are presented in the final two sections of this TM.

Ali, H. K.↗