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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 19 records
All sourcesNASA NTRSDOE OSTI
NASA NTRS1980

ac propulsion system for an electric vehicle

It is pointed out that dc drives will be the logical choice for current production electric vehicles (EV). However, by the mid-80's, there is a good chance that the price and reliability of suitable high-power semiconductors will allow for a competitive ac system. The driving force behind the ac approach is the induction motor, which has specific advantages relative to a dc shunt or series traction motor. These advantages would be an important factor in the case of a vehicle for which low maintenance characteristics are of primary importance. A description of an EV ac propulsion system is provided, taking into account the logic controller, the inverter, the motor, and a two-speed transmission-differential-axle assembly. The main barrier to the employment of the considered propulsion system in EV is not any technical problem, but inverter transistor cost.

Geppert, S.↗
NASA NTRS1980

The performance and efficiency of four motor/controller/battery systems for the simpler electric vehicles

A test and analysis program performed on four complete propulsion systems for an urban electric vehicle (EV) is described and results given. A dc series motor and a permanent magnet (PM) motor were tested, each powered by an EV battery pack and controlled by (1) a series/parallel voltage-switching (V-switch) system; and (2) a system using a pulse width modulation, 400 Hz transistorized chopper. Dynamometer tests were first performed, followed by eV performance predictions and data correlating road tests. During dynamometer tests using chopper control; current, voltage, and power were measured on both the battery and motor sides of the chopper, using three types of instrumentation. Conventional dc instruments provided adequate accuracy for eV power and energy measurements, when used on the battery side of the controller. When using the chopper controller, the addition of a small choke inductor improved system efficiency in the lower duty cycle range (some 8% increase at 50% duty cycle) with both types of motors. Overall system efficiency rankings during road tests were: (1) series motor with V-switch; (2) PM motor with V-switch; (3) series motor with chopper; and (4) PM motor with chopper. Chopper control of the eV was smoother and required less driver skill than V-switch control.

Shipps, P. R.↗
NASA NTRS2003

The Electric Propulsion Interactions Code (EPIC): A Member of the NASA Space Environment and Effects Program (SEE) Toolset

Science Applications International Corporation is currently developing the Electric Propulsion Interactions Code, EPIC, as part of a project sponsored by the Space Environments and Effects Program at NASA Marshall Space Flight Center. Now in its second year of development, EPIC is an interactive computer toolset that allows the construction of a 3-D spacecraft model, and the assessment of a variety of interactions between its subsystems and the plume from an electric thruster. This paper reports on the progress of EPZC including the recently added ability to exchange results the NASA Charging Analyzer Program, Nascap-2k. The capability greatly enhances EPIC's range of applicability. Expansion of the toolset's various physics models proceeds in parallel with the overall development of the software. Also presented are recent upgrades of the elastic scattering algorithm in the electric propulsion Plume Tool. These upgrades are motivated by the need to assess the effects of elastically scattered ions on the SIC for ion beam energies that exceed loo0 eV. Such energy levels are expected in future high-power (>10 kW) ion propulsion systems empowered by nuclear sources.

Mikellides, Ioannis G.↗
NASA NTRS2002

Magnetohydrodynamic Augmented Propulsion Experiment

A fundamental obstacle to routine space access is the specific energy limitations associated with chemical fuels. In the case of vertical take-off, the high thrust needed for vertical liftoff and acceleration to orbit translates into power levels in the 10 GW range. Furthermore, useful payload mass fractions are possible only if the exhaust particle energy (i.e., exhaust velocity) is much greater than that available with traditional chemical propulsion. The electronic binding energy released by the best chemical reactions (e.g., LOX/LH2 for example, is less than 2 eV per product molecule (approx. 1.8 eV per H2O molecule), which translates into particle velocities less than 5 km/s. Useful payload fractions, however, will require exhaust velocities exceeding 15 km/s (i.e., particle energies greater than 20 eV). As an added challenge, the envisioned hypothetical RLV (reusable launch vehicle) should accomplish these amazing performance feats while providing relatively low acceleration levels to orbit (2-3g maximum). From such fundamental considerations, it is painfully obvious that planned and current RLV solutions based on chemical fuels alone represent only a temporary solution and can only result in minor gains, at best. What is truly needed is a revolutionary approach that will dramatically reduce the amount of fuel and size of the launch vehicle. This implies the need for new compact high-power energy sources as well as advanced accelerator technologies for increasing engine exhaust velocity. Electromagnetic acceleration techniques are of immense interest since they can be used to circumvent the thermal limits associated with conventional propulsion systems. This paper describes the Magnetohydrodynamic Augmented Propulsion Experiment (MAPX) being undertaken at NASA Marshall Space Flight Center (MSFC). In this experiment, a 1-MW arc heater is being used as a feeder for a 1-MW magnetohydrodynamic (MHD) accelerator. The purpose of the experiment is to demonstrate that an MHD accelerator can be an effective augmentation system for increasing engine exhaust velocity. More specifically, the experiment is intended to show that electromagnetic effects are effective at producing flow acceleration whereas electrothermal effects do not cause unacceptable heating of the working fluid. The MHD accelerator was designed as an externally diagonalized segmented Faraday channel, which will be inserted into an existing 2-tesla electromagnet. This allows the external power to be connected through two terminals thereby minimizing the complexity and cost associated with powering each segment independently. The design of the accelerator and other components in the flow path has been completed and fabrication activities are underway. This paper provides a full description of MAPX including performance analysis, design, and test plans, and current status.

Litchford, Ron J.↗
NASA NTRS2009

Integrity of the Plasma Magnetic Nozzle

This report examines the physics governing certain aspects of plasma propellant flow through a magnetic nozzle, specifically the integrity of the interface between the plasma and the nozzle s magnetic field. The injection of 100s of eV plasma into a magnetic flux nozzle that converts thermal energy into directed thrust is fundamental to enabling 10 000s of seconds specific impulse and 10s of kW/kg specific power piloted interplanetary propulsion. An expression for the initial thickness of the interface is derived and found to be approx.10(exp -2) m. An algorithm is reviewed and applied to compare classical resistivity to gradient-driven microturbulent (anomalous) resistivity, in terms of the spatial rate and time integral of resistive interface broadening, which can then be related to the geometry of the nozzle. An algorithm characterizing plasma temperature, density, and velocity dependencies is derived and found to be comparable to classical resistivity at local plasma temperatures of approx. 200 eV. Macroscopic flute-mode instabilities in regions of "adverse magnetic curvature" are discussed; a growth rate formula is derived and found to be one to two e-foldings of the most unstable Rayleigh-Taylor (RT) mode. After establishing the necessity of incorporating the Hall effect into Ohm s law (allowing full Hall current to flow and concomitant plasma rotation), a critical nozzle length expression is derived in which the interface thickness is limited to about 1 ion gyroradius.

Gerwin, Richard A.↗
NASA NTRS1980

The adsorption of cesium on lanthanum hexaboride surfaces

The adsorption/desorption characteristics of cesium on clean and oxygen-covered LaB6 (100) surfaces were studied using various surface analysis techniques. On the initially clean surface (phi = 2.77 eV), adsorption produces a minimum work function of 1.96 eV and a saturation work function of 2.07 eV. For the oxygen-saturated surface, the cesium adsorption curve shows no minimum, the lowest work function (1.35 eV) occurring at cesium saturation. The utility of LaB6 (100) as an emitter or collector in thermionic converter applications (particularly for space nuclear electric propulsion) is evaluated.

Davis, P. R.↗
NASA NTRS2003

Simulation of Mini-Magnetospheric Plasma Propulsion (M2P2) Interacting with an External Plasma Wind

Substantial progress has been made over the last year in the development of the laboratory Mini-Magnetospheric Plasma Propulsion (M2P2) prototype. The laboratory testing has shown that that the plasma can be produced at high neutral gas efficiency, at high temperatures (a few tens of eV) with excellent confinement up to the point where chamber wall interactions dominate the physics. This paper investigates the performance of the prototype as it is opposed by an external plasma acting as a surrogate for the solar wind. The experiments were performed in 5ft diameter by 6ft long vacuum chamber at the University of Washington. The solar wind source comprised of a 33 kWe arc jet attached to a 200 kWe inductively generated plasma source. The dual plasma sources allow the interaction to be studied for different power levels, shot duration and production method. It is shown that plasma from the solar wind source (SWS) is able to penetrate the field of the M2P2 magnetic when no plasma is present. With operation of the M2P2 plasma source at only 1.5 kWe, the penetration of the SWS even at the highest power of operation at 200 kWe is stopped. This deflection is shown to be greatly enhanced over that produced by the magnet alone. In addition it is shown that with the presence of the SWS, M2P2 is able to produce enhanced magnetized plasma production out to at least 10 magnet radii where the field strength is only marginally greater than the terrestrial field. The results are consistent with the initial predictions that kWe M2P2 systems would be able to deflect several hundred kWe plasma winds to produce enhanced propulsion for a spacecraft.

Winglee, R. M.↗
NASA NTRS2004

Compact Plasma Accelerator

A plasma accelerator has been conceived for both material-processing and spacecraft-propulsion applications. This accelerator generates and accelerates ions within a very small volume. Because of its compactness, this accelerator could be nearly ideal for primary or station-keeping propulsion for spacecraft having masses between 1 and 20 kg. Because this accelerator is designed to generate beams of ions having energies between 50 and 200 eV, it could also be used for surface modification or activation of thin films.

Foster, John E.↗
NASA NTRS2002

Test and Analysis Capabilities of the Space Environment Effects Team at Marshall Space Flight Center

Marshall Space Flight Center has developed world-class space environmental effects testing facilities to simulate the space environment. The combined environmental effects test system exposes temperature-controlled samples to simultaneous protons, high- and low-energy electrons, vacuum ultraviolet (VUV) radiation, and near-ultraviolet (NUV) radiation. Separate chambers for studying the effects of NUV and VUV at elevated temperatures are also available. The Atomic Oxygen Beam Facility exposes samples to atomic oxygen of 5 eV energy to simulate low-Earth orbit (LEO). The LEO space plasma simulators are used to study current collection to biased spacecraft surfaces, arcing from insulators and electrical conductivity of materials. Plasma propulsion techniques are analyzed using the Marshall magnetic mirror system. The micro light gas gun simulates micrometeoroid and space debris impacts. Candidate materials and hardware for spacecraft can be evaluated for durability in the space environment with a variety of analytical techniques. Mass, solar absorptance, infrared emittance, transmission, reflectance, bidirectional reflectance distribution function, and surface morphology characterization can be performed. The data from the space environmental effects testing facilities, combined with analytical results from flight experiments, enable the Environmental Effects Group to determine optimum materials for use on spacecraft.

Finckenor, M. M.↗
NASA NTRS1962

ELECTRIC PROPULSION- A NEW TECHNOLOGY

Within the past 3 years, electric thrust devices have passed from the first exploratory feasibility experiments to the flight-engineered hardware phase. This rapid surge of development has been punctuated by equally rapid improvements in the performance figures of electrical thrust devices. The most spectacular advances were made in ion engine technology. Total energy expenditure per ion pair has gone from 35,000 ev in 1958 to about 600 ev at the present time leading to power utilization efficiencies of up to 80 per cent, depending upon the specific impulse. The percentage of the particles intercepting the accelerating and focusing electrodes has been reduced from more than 50 per cent in 1958 to less than 0.01 per cent at the present time. Great improvements in ion current per unit area have also been realized, leading to greater thrust "pressures."

PLASMA ENGINE↗
NASA NTRS1977

Cross sections for charge transfer between mercury ions and other metals

Cross sections for charge transfer between several ions and metals of interest to the NASA electro propulsion program have been measured. Specifically, the ions considered were Hg(+), Xe(+) and Cs(+) and the metals Mo, Fe, Al, Ti, Ta, and C. Measurements were made in the energy regime from 1 to 5,000 eV. In general, the cross sections for charge transfer were found to be less than 10 to the minus 15 power sq cm for most processes over the total energy range. Exceptions are Hg(+) in collision with Ti and Ta. The results obtained for each reaction are given in both graphical and numerical form in the text. For quick reference, the data at several ion velocities are condensed into one table given in the summary.

Vroom, D. A.↗
NASA NTRS2018

Overview of Phobos/Deimos Regolith Ion Sample Mission (PRISM) Concept

Far more definitive information on composition is required to resolve the question of origin for the Martian moons Phobos and Deimos. Current infrared spectra of the objects are inconclusive due to the lack of strong diagnostic features.Definitive compositional measurements of Phobos could be obtained using in-situ X-ray, gamma-ray, or neutronspectroscopy or collecting and returning samples to Earth for analysis. We have proposed, in lieu of those methods, toderive Phobos and Deimos compositional data from secondary ion mass spectrometry (SIMS) measurements by calibratingthe instrument to elemental abundance measurements made for known samples in the laboratory. We describe thePhobos/Deimos Regolith Ion Sample Mission (PRISM) concept here. PRISM utilizes a high-resolution TOF plasma composition analyzer to make SIMS measurements by observing the sputtered species from various locations of the moons' surfaces. In general, the SIMS technique and ion mass spectrometers complement and expand quadrupole mass spectrometer measurements by collecting ions that have been energized to higher energies, 50-100 eV, and making measurements at very low densities and pressures. Furthermore, because the TOF technique accepts all masses all the time,it obtains continuous measurements and does not require stepping through masses. The instrument would draw less than10 W and weigh less than 5 kg. The spacecraft, nominally a radiation-hardened 12U CubeSat, would use a low-thrust SolarElectric Propulsion system to send it on a two-year journey to Mars, where it would co-orbit with Deimos and then Phobos

Clark, Pamela↗
NASA NTRS2018

Overview of Phobos/Deimos Regolith Ion Sample Mission (PRISM) Concept

Far more definitive information on composition is required to resolve the question of origin for the Martian moons Phobos and Deimos. Current infrared spectra of the objects are inconclusive due to the lack of strong diagnostic features. Definitive compositional measurements of Phobos could be obtained using in-situ X-ray, gamma-ray, or neutron spectroscopy or collecting and returning samples to Earth for analysis. We have proposed, in lieu of those methods, to derive Phobos and Deimos compositional data from secondary ion mass spectrometry (SIMS) measurements by calibrating the instrument to elemental abundance measurements made for known samples in the laboratory. We describe the Phobos/Deimos Regolith Ion Sample Mission (PRISM) concept here. PRISM utilizes a high-resolution TOF plasma composition analyzer to make SIMS measurements by observing the sputtered species from various locations of the moons' surfaces. In general, the SIMS technique and ion mass spectrometers complement and expand quadrupole mass spectrometer measurements by collecting ions that have been energized to higher energies, 50-100 eV, and making measurements at very low densities and pressures. Furthermore, because the TOF technique accepts all masses all the time, it obtains continuous measurements and does not require stepping through masses. The instrument would draw less than 10 W and weigh less than 5 kg. The spacecraft, nominally a radiation-hardened 12U CubeSat, would use a low-thrust Solar Electric Propulsion system to send it on a two-year journey to Mars, where it would co-orbit with Deimos and then Phobos at distances as low as 27 km.

Dibraccio, Gina↗
NASA NTRS1989

An RF heated tandem mirror plasma propulsion study

Experimental results on a tandem mirror hybrid plume rocket involving a three-stage system of plasma injection, heating, and subsequent injection through a magnetic nozzle are presented. In the experiments, a plasma is created by breaking down the gas with electron cyclotron resonance heating at 2 kW in the central cell, and the ion species is then heated to high temperatures with ion cyclotron resonance heating at 10 kW in the end cell. A Langmuir probe measured an electron density of 2.5 x 10 to the 16th/cu m and a temperature of 100 eV in the central cell and an ion density of 1.25 x 10 to the 17th/cu m and a temperature of 500 eV in the end cell.

Yang, T. F.↗
NASA NTRS2015

Near-Surface Plasma Characterization of the 12.5-kW NASA TDU1 Hall Thruster

To advance the state-of-the-art in Hall thruster technology, NASA is developing a 12.5-kW, high-specific-impulse, high-throughput thruster for the Solar Electric Propulsion Technology Demonstration Mission. In order to meet the demanding lifetime requirements of potential missions such as the Asteroid Redirect Robotic Mission, magnetic shielding was incorporated into the thruster design. Two units of the resulting thruster, called the Hall Effect Rocket with Magnetic Shielding (HERMeS), were fabricated and are presently being characterized. The first of these units, designated the Technology Development Unit 1 (TDU1), has undergone extensive performance and thermal characterization at NASA Glenn Research Center. A preliminary lifetime assessment was conducted by characterizing the degree of magnetic shielding within the thruster. This characterization was accomplished by placing eight flush-mounted Langmuir probes within each discharge channel wall and measuring the local plasma potential and electron temperature at various axial locations. Measured properties indicate a high degree of magnetic shielding across the throttle table, with plasma potential variations along each channel wall being less than or equal to 5 eV and electron temperatures being maintained at less than or equal to 5 eV, even at 800 V discharge voltage near the thruster exit plane. These properties indicate that ion impact energies within the HERMeS will not exceed 26 eV, which is below the expected sputtering threshold energy for boron nitride. Parametric studies that varied the facility backpressure and magnetic field strength at 300 V, 9.4 kW, illustrate that the plasma potential and electron temperature are insensitive to these parameters, with shielding being maintained at facility pressures 3X higher and magnetic field strengths 2.5X higher than nominal conditions. Overall, the preliminary lifetime assessment indicates a high degree of shielding within the HERMeS TDU1, effectively mitigating discharge channel erosion as a life-limiting mechanism.

Lifetime↗
NASA NTRS2004

Magnetic-Nozzle Studies for Fusion Propulsion Applications: Gigawatt Plasma Source Operation and Magnetic Nozzle Analysis

This project has been a multiyear effort to assess the feasibility of a key process inherent to virtually all fusion propulsion concepts: the expansion of a fusion-grade plasma through a diverging magnetic field. Current fusion energy research touches on this process only indirectly through studies of plasma divertors designed to remove the fusion products from a reactor. This project was aimed at directly addressing propulsion system issues, without the expense of constructing a fusion reactor. Instead, the program designed, constructed, and operated a facility suitable for simulating fusion reactor grade edge plasmas, and to examine their expansion in an expanding magnetic nozzle. The approach was to create and accelerate a dense (up to l0(exp 20)/m) plasma, stagnate it in a converging magnetic field to convert kinetic energy to thermal energy, and examine the subsequent expansion of the hot (100's eV) plasma in a subsequent magnetic nozzle. Throughout the project, there has been a parallel effort between theoretical and numerical design and modelling of the experiment and the experiment itself. In particular, the MACH2 code was used to design and predict the performance of the magnetoplasmadynamic (MPD) plasma accelerator, and to design and predict the design and expected behavior for the magnetic field coils that could be added later. Progress to date includes the theoretical accelerator design and construction, development of the power and vacuum systems to accommodate the powers and mass flow rates of interest to out research, operation of the accelerator and comparison to theoretical predictions, and computational analysis of future magnetic field coils and the expected performance of an integrated source-nozzle experiment.

Gilland, James H.↗
NASA NTRS2004

Discharge Chamber Performance of the NEXIS Ion Thruster

The Nuclear-Electric Xenon Ion System (NEXIS) thruster was designed to produce greater than or equal to 70% efficiency at ISPs in excess of 6500 sec and total power levels in excess of 15 kW. In order to achieve this performance, the thruster requires a large area plasma generator capable of high propellant utilimtion efficiency and low discharge loss while producing a very flat, uniform beam profile. Fortunately, larger thrusters can be made more uniform and efficient due to the higher volume to surface ratio, provided that the magnetic cusp confinement is designed properly and the thruster length to diameter ratio is adequate. This paper describes the discharge chamber performance of the NEXIS Laboratory Model (LM) thruster. The LM discharge chamber is 65 cm in diameter at the grid plane and uses 6 ring-cusps to provide magnetic confinement of the plasma. The thruster was tested with flat carbon-carbon composite grids with the hole pattern masked to 57 cm in diameter and a conventional Type-B "1/2" diameter hollow cathode. During the preliminary "discharge only" tests, the LM thruster demonstrated profile factors of 0.84 and a discharge loss of about 160 eV/ion at 25 V discharge voltage and over 90% propellant utilization efficiency in simulated beam extraction experiments at 3.9 A of beam current. Analysis of the data from these tests used the discharge-only model developed by Brophy. Subsequent beam extraction experiments validated the key variables used in the model to predict the performance from the discharge-only data, and demonstrated 3.9 A of beam current at over 90% propellant utilization efficiency with a flatness parameter of better than 0.8 and a discharge loss of about 185 eV/ion. The slightly higher discharge loss measured during beam extractions was found to be due to a lower screen transparency in the as-manufactured LM grid set. Plasma measurements with a scanning probe internal to the thruster near the screen grid showed plasma densities over l x 10(exp 11) per cubic centimeter and electron temperatures of 3.5 to 5.5 eV depending on the operation parameters. The performance of the NEXIS discharge chamber contributed to the over 78% thruster efficiency measured during beam extraction at 7500 sec ISP and 25 kW of power, and over 81% thruster efficiency measured at 8500 sec ISP.

plasma generation↗
NASA NTRS2004

High Power Light Gas Helicon Plasma Source For VASMIR

The VASIMR space propulsion development effort relies on a high power (greater than 10kW) helicon source to produce a dense flowing plasma (H, D and He) target for ion cyclotron resonance (ICR) acceleration of the ions. Subsequent expansion in an expanding magnetic field (magnetic nozzle) converts ion lunetic energy to directed momentum. This plasma source must have critical features to enable an effective propulsion device. First, it must ionize most of the input neutral flux of gas, thus producing a plasma stream with a high degree of ionization for application of ICR power. This avoids propellant waste and potential power losses due to charge exchange. Next, the plasma stream must flow into a region of high magnetic field (approximately 0.5 T) for efficient ICR acceleration. Third, the ratio of input power to plasma flux must be low, providing an energy per ion-electron pair approaching 100 eV. Lastly, the source must be robust and capable of very long life-times (years). In our helicon experiment (VX-10) we have measured a ratio of input gas to plasma flux near 100%. The plasma flows from the helicon region (B approximately 0.1 T) into a region with a peak magnetic field of 0.8 T. The energy input per ion-electron pair has been measured at 300 plus or minus 100 eV. Recent results at Oak Ridge National Laboratory (ORNL) show an enhanced efficiency mode of operation with a high power density, over 5 kW in a 5 cm diameter tube. Our helicon is presently 9 cm in diameter and operates up to 3.5 kW of input power. An upgrade to a power level of 10 kW is underway. Much of our recent work has been with a Boswell double-saddle antenna design. We are also converting the antenna design to a helical type. With these modifications, we anticipate an improvement in the ionization efficiency. This paper presents the results from scaling the helicon in the VX-10 device from 3.5 to 10 kW. We also compare the operation with a double-saddle to a helical antenna design. Finally, we discuss modeling of these configurations using ORNL's EMIR code.

Squire, J. P.↗