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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 901 records · Page 50
All sourcesNASA NTRSDOE OSTI
NASA NTRS2014

Direct Energy Conversion for Nuclear Propulsion at Low Specific Mass

The project will continue the FY13 JSC IR&D (October-2012 to September-2013) effort in Travelling Wave Direct Energy Conversion (TWDEC) in order to demonstrate its potential as the core of a high potential, game-changing, in-space propulsion technology. The TWDEC concept converts particle beam energy into radio frequency (RF) alternating current electrical power, such as can be used to heat the propellant in a plasma thruster. In a more advanced concept (explored in the Phase 1 NIAC project), the TWDEC could also be utilized to condition the particle beam such that it may transfer directed kinetic energy to a target propellant plasma for the purpose of increasing thrust and optimizing the specific impulse. The overall scope of the FY13 first-year effort was to build on both the 2012 Phase 1 NIAC research and the analysis and test results produced by Japanese researchers over the past twenty years to assess the potential for spacecraft propulsion applications. The primary objective of the FY13 effort was to create particle-in-cell computer simulations of a TWDEC. Other objectives included construction of a breadboard TWDEC test article, preliminary test calibration of the simulations, and construction of first order power system models to feed into mission architecture analyses with COPERNICUS tools. Due to funding cuts resulting from the FY13 sequestration, only the computer simulations and assembly of the breadboard test article were completed. The simulations, however, are of unprecedented flexibility and precision and were presented at the 2013 AIAA Joint Propulsion Conference. Also, the assembled test article will provide an ion current density two orders of magnitude above that available in previous Japanese experiments, thus enabling the first direct measurements of power generation from a TWDEC for FY14. The proposed FY14 effort will use the test article for experimental validation of the computer simulations and thus complete to a greater fidelity the mission analysis products originally conceived for FY13.

Scott, John H.↗
NASA NTRS1979

Reduced power processor requirements for the 30-cm diameter HG ion thruster

The characteristics of power processors strongly impact the overall performance and cost of electric propulsion systems. A program was initiated to evaluate simplifications of the thruster-power processor interface requirements. The power processor requirements are mission dependent with major differences arising for those missions which require a nearly constant thruster operating point (typical of geocentric and some inbound planetary missions) and those requiring operation over a large range of input power (such as outbound planetary missions). This paper describes the results of tests which have indicated that as many as seven of the twelve power supplies may be eliminated from the present Functional Model Power Processor used with 30-cm diameter Hg ion thrusters.

Rawlin, V. K.↗
NASA NTRS1976

A 10,000 hour endurance test of a 700 series 30 cm engineering model thruster

This paper describes a 10,000 hour test of the first engineering model of the 30-cm ion thruster which is expected to form the basis of future prime propulsion systems. This test demonstrated the lifetime capability of such critical components as cathodes, vaporizers, isolators, and optics. The results of the test are described including the conclusions derived from an extensive post-test analysis of the thruster. This test did identify a major life limiting problem - discharge chamber erosion. The modifications which have been incorporated in the 900 series 30-cm ion thrusters to minimize this problem are described.

Collett, C. R.↗
NASA NTRS2007

Status of the Space-Rated Lithium-Ion Battery Advanced Development Project in Support of the Exploration Vision

The NASA Glenn Research Center (GRC), along with the Goddard Space Flight Center (GSFC), Jet Propulsion Laboratory (JPL), Johnson Space Center (JSC), Marshall Space Flight Center (MSFC), and industry partners, is leading a space-rated lithium-ion advanced development battery effort to support the vision for Exploration. This effort addresses the lithium-ion battery portion of the Energy Storage Project under the Exploration Technology Development Program. Key discussions focus on the lithium-ion cell component development activities, a common lithium-ion battery module, test and demonstration of charge/discharge cycle life performance and safety characterization. A review of the space-rated lithium-ion battery project will be presented highlighting the technical accomplishments during the past year.

Miller, Thomas↗
NASA NTRS1985

Microwave electric propulsion for orbit transfer applications

A Microwave Electric Propulsion (MEP) concept is developed for an unmanned Orbit Transfer Vehicle (OTV). The concept is based on the use of beamed microwave power and on an extrapolation of ion thruster technology. Beamed microwave power transmission is discussed in terms of its relationship to spacecraft propulsion. Characteristics of an MEP OTV are determined and performance is evaluated by comparison to a variety of alternative propulsion systems for the completion of a ten-year mission model.

Sercel, J. C.↗
NASA NTRS2019

Far Field Plume Distribution and Divergence for NEXT: DART Mission

In support of the Double Asteroid Redirection Test (DART) mission, laboratory measurements were made on the NEXT ion engine, which will be used for the spacecraft's in-space propulsion [1]. This study revisits a small range of mission-specific 2.7A throttle levels to understand the effect of in-flight flow rate variability, investigate intermediate throttle conditions, and improve measurement methodology. This paper specifically examines the far-field plume divergence and backflow ion flux distribution of the NEXT, while a companion paper examines the charge state distributions.

Throttle Levels↗
NASA NTRS2002

Extending Ion Engine Technology to NEXT and Beyond

Extending ion engine technology beyond the current state-of-the art primary interplanetary electric propulsion system, the 2.3-kW NASA Solar Electric Propulsion Technology and Applications Readiness (NSTAR) system, will require thrusters with improved propellant throughput and total impulse capability. Many of the design choices that culminated in the NSTAR thrusters must be revisited, and their application to next generation ion engine technology must be evaluated. The concept of derating, which was successfully employed in NSTAR, has been applied to the 40 cm NASA Evolutionary Xenon Thruster (NEXT) currently under development at NASA Glenn Research Center (GRC). At 5-kW, NEXT operates with the same average beam current density as NSTAR, and at 10-kW, the peak beam current density is only ten percent greater than NSTAR. The result is that similar Ion optics technology is expected to yield comparable lifetime. Thick-accelerator- grid ion optics are also being tested to realize additional lifetime benefits. A 40-A discharge cathode is being developed for NEXT based on scaling the NSTAR design. Nevertheless, the experiences of the NSTAR ground tests and the thruster on the Deep Space One spacecraft indicate that the discharge cathode wear must be studied experimentally and theoretically to ensure that it meets the lifetime requirements. Although NEXT is in its infancy, investigations have already begun to examine possible modifications to engine design for even higher-power and higher-specific impulse engines. Ion optics using alternate materials such as titanium, graphite, or carbon-carbon composite are currently being investigated due to their low sputter yields at high voltage. To avoid the difficulties encountered using electrodes at high-currents, the use of a microwave-based ion thruster is under investigation for potential high-power ion thruster systems requiring long lifetimes. Additionally, alternative propellants are being considered for applications requiring high-specific impulse (>> 5000 s) and extremely long-life (>> 15,000 hr). Testing requirements make condensable propellants attractive for high-power engines. Although the NSTAR ion engine demonstrated the flight maturity of ion thruster technology, many challenges remain for the development of thrusters with improved propellant throughput and power handling capabilities.

Domonkos, Matthew T.↗
NASA NTRS1973

The advantages of the high voltage solar array for electric propulsion

The high voltage solar array (HVSA) offers improvements in efficiency, weight, and reliability for the electric propulsion power system. The basic HVSA technology involves designing the solar array to deliver power in the form required by the ion thruster. This paper delves into conventional power processes and problems associated with ion thruster operation using SERT II experience for examples. In this light, the advantages of the HVSA concept for electric propulsion are presented. Tests conducted operating the SERT II thruster system in conjunction with HVSA are discussed. Thruster operation was observed to be normal and in some respects improved.

Sater, B. L.↗
NASA NTRS1979

Hg ion thruster component testing

Electron bombardment thrusters, under development to provide both auxiliary and primary propulsion functions for a large variety of space missions are tested. Thruster design verification which requires life tests of durations of the order of the time anticipated in space applications, are discussed. The life time and reliability of an electron bombardment thruster is dependent upon the performance of several critical components including cathodes, vaporizers, and isolators. The performances of the cathode, vaporizer, and propellant isolaters during fatigue analyses are examined.

Mantenieks, M. A.↗
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 NTRS1981

Performance of a magnetic multipole line-cusp argon ion thruster

A 17 cm diameter line cusp ion thruster was evaluated with inert gases which are candidate propellants for on orbit and orbit transfer propulsion functions for Large Space Systems. A semiempirical relationship was generated to predict thruster beam current in terms of plasma parameters which would allow initial thruster optimization without ion extraction and the associated large vacuum facilities. The sensitivity of performance to changes in discharge electrode configurations and magnetic circuit was evaluated and is presented. After final optimization a propellant utilization efficiency of 0.9 at a discharge chamber power expenditure of about 260 w per beam ampere was obtained. These performance parameters are the highest yet achieved with argon propellant.

Sovey, J. S.↗
NASA NTRS1978

NASA Electric Propulsion Program

The described NASA program is to provide the technology base to meet the U.S. future needs in the area of low thrust, high specific impulse propulsion. An electric propulsion system will be needed to augment the Shuttle-IUS Space Transportation System for space missions involving the exploration of the solar system. Plans exist in this connection for comet and asteroid rendezvous, a Saturn and a Mercury orbiter, and a Mars surface sample return. An analysis has shown that electric propulsion can be used to obtain larger payloads, shorter trip times, more mission flexibility, and less complicated trajectories. In addition, electric propulsion has also potential uses for earth orbital missions. Attention is given to the auxiliary propulsion technology program which is to develop a one millipound electron bombardment thrust system, a baseline program concerned with a 3 kilowatt thrust system, ion thruster research, magnetoplasmadynamic thruster research and evaluation, and the use of a mass driver as a reaction mass propulsion system.

Hudson, W. R.↗
NASA NTRS1992

Micro Electric Propulsion Feasibility

Miniature, 50 kg class, strategic satellites intended for extended deployment in space require an on-board propulsion capability to perform needed attitude control adjustments and drag compensation maneuvers. Even on such very small spacecraft, these orbit maintenance functions can be significant and result in a substantial propellant mass requirement. Development of advanced propulsion technology could reduce this propellant mass significantly, and thereby maximize the payload capability of these spacecraft. In addition, spacecraft maneuverability could be enhanced and/or multi-year mission lifetimes realized. These benefits cut spacecraft replacement costs, and reduce services needed to maintain the launch vehicles. For SDIO brilliant pebble spacecraft, a miniaturized hydrazine propulsion system provides both boost and divert thrust control. This type of propulsion system is highly integrated and is capable of delivering large thrust levels for short time periods. However, orbit maintenance functions such as drag make-up require only very small velocity corrections. Using the boost and/or divert thrusters for these small corrections exposes this highly integrated propulsion system to continuous on/off cycling and thereby increases the risk of system failure. Furthermore, since drag compensation velocity corrections would be orders of magnitude less than these thrusters were designed to deliver, their effective specific impulse would be expected to be lower when operated at very short pulse lengths. The net result of these effects would be a significant depletion of the on-board hydrazine propellant supply throughout the mission, and a reduced propulsion system reliability, both of which would degrade the interceptors usefulness. In addition to SDIO brilliant pebble spacecraft, comparably small spacecraft can be anticipated for other future strategic defense applications such as surveillance and communication. For such spacecraft, high capability and reliability, minimal detectability and low cost are requirements. All these miniature spacecraft share a common characteristic: because of their on-board electronic equipment they have, by design, solar order 50-100 W. In a relative sense, such spacecraft are power rich when compared to other larger spacecraft. This power rich situation is offset by very tight mass budgets, which make reductions in propellant mass requirements a key issue in meeting overall spacecraft minimum mass goals. In principle, power rich and propellant poor brilliant pebbles class spacecraft can benefit from using high specific impulse electric propulsion to reduce chemical propellant mass requirements. However, at power levels of order 50 W, arcjets cannot be made to function, ion thrusters are too complex and heavy and resistojets have too low a specific impulse. Recognizing these capability limitations in existing electric propulsion technology, the SDIO/IST sponsored the Phase I SBIR Micro Electric Propulsion (MEP) thruster study described in this report. The objective of this study was to examine the feasibility of developing a very simple, low mass and small volume, electric thruster for operation on hydrazine at less than 100 W of input power. The feasibility of developing such a MEP thruster was successfully demonstrated by EPL by the discovery of a novel plasma acceleration process. The sections in this report summarize the approach, test results and major accomplishments of this proof-of-concept program.

Aston, Graeme↗
NASA NTRS1999

Advanced Electric Propulsion for Space Solar Power Satellites

The sun tower concept of collecting solar energy in space and beaming it down for commercial use will require very affordable in-space as well as earth-to-orbit transportation. Advanced electric propulsion using a 200 kW power and propulsion system added to the sun tower nodes can provide a factor of two reduction in the required number of launch vehicles when compared to in-space cryogenic chemical systems. In addition, the total time required to launch and deliver the complete sun tower system is of the same order of magnitude using high power electric propulsion or cryogenic chemical propulsion: around one year. Advanced electric propulsion can also be used to minimize the stationkeeping propulsion system mass for this unique space platform. 50 to 100 kW class Hall, ion, magnetoplasmadynamic, and pulsed inductive thrusters are compared. High power Hall thruster technology provides the best mix of launches saved and shortest ground to Geosynchronous Earth Orbital Environment (GEO) delivery time of all the systems, including chemical. More detailed studies comparing launch vehicle costs, transfer operations costs, and propulsion system costs and complexities must be made to down-select a technology. The concept of adding electric propulsion to the sun tower nodes was compared to a concept using re-useable electric propulsion tugs for Low Earth Orbital Environment (LEO) to GEO transfer. While the tug concept would reduce the total number of required propulsion systems, more launchers and notably longer LEO to GEO and complete sun tower ground to GEO times would be required. The tugs would also need more complex, longer life propulsion systems and the ability to dock with sun tower nodes.

Oleson, Steve↗
NASA NTRS1989

A brief review of JPL's electric propulsion technology activities

Near-term objectives and recent technological progress of JPL's electric propulsion program are discussed. Particular attention is given to accomplishments for ion, magnetoplasmadynamic (MPD), electron-cyclotron resonance (ECR), and arcjet thrusters. Xenon ion thruster erosion tests indicate a 15-fold reduction in tantalum baffle erosion when nitrogen is added to the xenon propellant and steady-state cylindrical MPD thruster tests at powers up to 72 kW show distinct self-constricted and diffuse discharge modes. An ECR thruster was operated at up to 7 kW with plasma acceleration at energies up to 7 kW; there was plasma acceleration at energies approaching 100 electron volts.

Barnett, John W.↗
NASA NTRS2002

NEMS for space and terrestrial applications

Nanoscale electromechanical systems (NEMS) techniques were used to manufacture a membrane used to supply singly charged un-fractured ions for mass spectrometry. The innovative nanostructure, developed at the Jet Propulsion Laboratory (JPL), is a major advance over micromachined 'volcano' field ionization devices and filaments or other energy-intensive devices, achieving nearly 100% ionization of gases and vapors without appreciable fracture.

Nanoscale electromechanical systems NEMS soft ioni↗
NASA NTRS2014

A Cubesat Asteroid Mission: Propulsion Trade-offs

A conceptual design was performed for a 6-U cubesat for a technology demonstration to be launched on the NASA Space Launch System (SLS) test launch EM-1, to be launched into a free-return translunar trajectory. The mission purpose was to demonstrate use of electric propulsion systems on a small satellite platform. The candidate objective chosen was a mission to visit a Near-Earth asteroid. Both asteroid fly-by and asteroid rendezvous missions were analyzed. Propulsion systems analyzed included cold-gas thruster systems, Hall and ion thrusters, incorporating either Xenon or Iodine propellant, and an electrospray thruster. The mission takes advantage of the ability of the SLS launch to place it into an initial trajectory of C3=0.

corrosion↗