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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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NASA NTRS1985

Shuttle electrical environment

Part of an AFGL payload flown on the STS-4 mission consisted of experiments to measure in-situ electric fields, electron densities, and vehicle charging. During this flight some 11 hours of data were acquired ranging from 5 minute snapshots up to continuous half-orbits. These experiments are described and results presented for such vehicle induced events as a main engine burn, thruster firings and water dumps in addition to undisturbed periods. The main characteristic of all the vehicle induced events is shown to be an enhancement in the low frequency noise (less than 2 kHz), in both the electrostatic and electron irregularity (delta N/N) spectra. The non-event results indicate that the electrostatic broadband emissions show a white noise characteristic in the low frequency range up to 2 kHz at an amplitude of 10 db above the shuttle design specification limit, falling below that limit above 10 kHz. The vehicle potential remained within the range of -3 to +1 volt throughout the flight which exhibits normal behavior for a satellite in a low equatorial orbit.

Smiddy, M.↗

NASA NTRS1973

Electrostatic Propulsion Beam Divergence Effects on Spacecraft Surfaces. Volume 2, Addendum 1: Ion Time-of-flight Determinations of Doubly to Singly Ionized Mercury Ion Ratios from a Mercury Electron Bombardment Discharge

The analysis of ion exhaust beam current flow for multiply charged ion species and the application to propellant utilization for the thruster are discussed. The ion engine in use in the experiments is a twenty centimeter diameter electromagnet electron bombardment engine. The experimental technique to determine the multiply charged ion abundance ratios using ion time of flight is described. An analytical treatment of the discharge action in producing various ion species has been carried out.

Sellen, J. M., Jr.↗

NASA NTRS2013

Azimuthal spoke propagation in Hall Effect Thrusters

Spokes are azimuthally propagating perturbations in the plasma discharge of Hall Effect Thrusters (HETs) that travel in the E × B direction and have been observed in many different systems. The propagation of azimuthal spokes are investigated in a 6 kW HET known as the H6 using ultra-fast imaging and azimuthally spaced probes. A spoke surface is a 2-D plot of azimuthal light intensity evolution over time calculated from 87,500 frames/s videos. The spoke velocity has been determined using three methods with similar results: manual fitting of diagonal lines on the spoke surface, linear cross-correlation between azimuthal locations and an approximated dispersion relation. The spoke velocity for three discharge voltages (300, 400 and 450 V) and three anode mass flow rates (14.7, 19.5 and 25.2 mg/s) yielded spoke velocities between 1500 and 2200 m/s across a range of normalized magnetic field settings. The spoke velocity was inversely dependent on magnetic field strength for low B-field settings and asymptoted at B-field higher values. The velocities and frequencies are compared to standard drifts and plasma waves such as E × B drift, electrostatic ion cyclotron, magnetosonic and various drift waves. The empirically approximated dispersion relation yielded a characteristic velocity that matched the ion acoustic speed for 5 eV electrons that exist in the near-anode and near-field plume regions of the discharge channel based on internal measurements. Thruster performance has been linked to operating mode where thrust-to-power is maximized when azimuthal spokes are present so investigating the underlying mechanism of spokes will benefit thruster operation.

Polk, James E.↗

NASA NTRS2013

Azimuthal Spoke Propagation in Hall Effect Thrusters

Spokes are azimuthally propagating perturbations in the plasma discharge of Hall Effect Thrusters (HETs) that travel in the E x B direction and have been observed in many different systems. The propagation of azimuthal spokes are investigated in a 6 kW HET known as the H6 using ultra-fast imaging and azimuthally spaced probes. A spoke surface is a 2-D plot of azimuthal light intensity evolution over time calculated from 87,500 frames/s videos. The spoke velocity has been determined using three methods with similar results: manual fitting of diagonal lines on the spoke surface, linear cross-correlation between azimuthal locations and an approximated dispersion relation. The spoke velocity for three discharge voltages (300, 400 and 450 V) and three anode mass flow rates (14.7, 19.5 and 25.2 mg/s) yielded spoke velocities between 1500 and 2200 m/s across a range of normalized magnetic field settings. The spoke velocity was inversely dependent on magnetic field strength for low B-field settings and asymptoted at B-field higher values. The velocities and frequencies are compared to standard drifts and plasma waves such as E x B drift, electrostatic ion cyclotron, magnetosonic and various drift waves. The empirically approximated dispersion relation yielded a characteristic velocity that matched the ion acoustic speed for 5 eV electrons that exist in the near-anode and near-field plume regions of the discharge channel based on internal measurements. Thruster performance has been linked to operating mode where thrust-to-power is maximized when azimuthal spokes are present so investigating the underlying mechanism of spokes will benefit thruster operation.

Cylindrical Hall Thruster (CHT)↗

NASA NTRS1987

Electromagnetic emission experiences using electric propulsion systems: A survey

As electric propulsion systems become ready to integrate with spacecraft systems, the impact of propulsion system radiated emissions are of significant interest. Radiated emissions from electromagnetic, electrostatic, and electrothermal systems have been characterized and results synopsized from the literature describing 21 space flight programs. Electromagnetic radiated emission results from ground tests and flight experiences are presented with particular attention paid to the performance of spacecraft subsystems and payloads during thruster operations. The impacts to transmission of radio frequency signals through plasma plumes are also reviewed.

Sovey, James S.↗

NASA NTRS1987

Electromagnetic emission experiences using electric propulsion systems - A survey

Sovey, James S.↗

NASA NTRS1990

Current-driven plasma acceleration versus current-driven energy dissipation. I - Wave stability theory

The dominant unstable electrostatic wave modes of an electromagnetically accelerated plasma are investigated. The study is the first part of a three-phase program aimed at characterizing the current-driven turbulent dissipation degrading the efficiency of Lorentz force plasma accelerators such as the MPD thruster. The analysis uses a kinetic theory that includes magnetic and thermal effects as well as those of an electron current transverse to the magnetic field and collisions, thus combining all the features of previous models. Analytical and numerical solutions allow a detailed description of threshold criteria, finite growth behavior, destabilization mechanisms and maximized-growth characteristics of the dominant unstable modes. The lower hybrid current-driven instability is implicated as dominant and was found to preserve its character in the collisional plasma regime.

Kelly, A. J.↗

NASA NTRS2010

Investigation of a Gallium MPD Thruster with an Ablating Cathode

Arc impedance, exhaust velocity, and plasma probe measurements are presented. The thruster is driven by a 50 microsecond pulse from a 6.2 milliohm pulse forming network, and gallium is supplied to the discharge by evaporation of the cathode. The arc voltage is found to vary linearly with the discharge current with an arc impedance of 6.5 milliohms. Electrostatic probes yield an exhaust velocity that is invariant with the discharge current and has a peak value of 20 kilometers per second, which is in reasonable agreement with the value (16 plus or minus 1 kilometer per second) calculated from the mass bit and discharge current data. Triple probe measurements yield on axis electron temperatures in the range of 0.8-3.8 eV, electron densities in the range of 1.6 x 10(exp 21) to 2.1 x 10(exp 22) per cubic meter, and a divergence half angle of 16 degrees. Measurements within the interelectrode region yield a peak magnetic field of 0.8 T, and the observed radial trends are consistent with an azimuthally symmetric current distribution. A cathode power balance model is coupled with an ablative heat conduction model predicting mass bit values that are within 20% of the experimental values.

Thomas, Robert E.↗

NASA NTRS1993

Low power arcjet system spacecraft impacts

Potential plume contamination of spacecraft surfaces was investigated by positioning spacecraft material samples relative to an arcjet thruster. Samples in the simulated solar array region were exposed to the cold gas arcjet plume for 40 hrs to address concerns about contamination by backstreaming diffusion pump oil. Except for one sample, no significant changes were measured in absorptance and emittance within experimental error. Concerns about surface property degradation due to electrostatic discharges led to the investigation of the discharge phenomenon of charged samples during arcjet ignition. Short duration exposure of charged samples demonstrated that potential differences are consistently and completely eliminated within the first second of exposure to a weakly ionized plume. The spark discharge mechanism was not the discharge phenomenon. The results suggest that the arcjet could act as a charge control device on spacecraft.

Pencil, Eric J.↗

NASA NTRS1979

Electromagnetic interference assessment of an ion drive electric propulsion system

The electromagnetic interference (EMI) form elements of an ion drive electric propulsion system was analyzed, and the effects of EMI interaction with a typical interplanetary spacecraft engineering and scientific subsystems were predicted. SEMCAP, a computerized electromagnetic compatibility assessment code, was used to analyze the impact of EMI noise sources on 65 engineering/telemetry circuits and 48 plasma wave and planetary radio astronomy channels measuring over the range of 100 Hz to 40 MHz in a spacecraft of the Voyager type; manual methods were used to evaluate electrostatics, magnetics, and communications effects. Results indicate that some conducted and radiated spectra are in excess of electromagnetic compatibility specification limits; direct design changes may be required for filtering and shielding of thrust system elements. The worst source of broadband radiated noise appears to be the power processor. The magnetic field necessary to thruster operation is equivalent to about 18 amp-sq m per amp of beam current at right angles to the axis caused by the neutralizer/plume loop.

Whittlesey, A. C.↗

NASA NTRS2004

Asymmetrical Capacitors for Propulsion

Asymmetrical Capacitor Thrusters have been proposed as a source of propulsion. For over eighty years, it has been known that a thrust results when a high voltage is placed across an asymmetrical capacitor, when that voltage causes a leakage current to flow. However, there is surprisingly little experimental or theoretical data explaining this effect. This paper reports on the results of tests of several Asymmetrical Capacitor Thrusters (ACTs). The thrust they produce has been measured for various voltages, polarities, and ground configurations and their radiation in the VHF range has been recorded. These tests were performed at atmospheric pressure and at various reduced pressures. A simple model for the thrust was developed. The model assumed the thrust was due to electrostatic forces on the leakage current flowing across the capacitor. It was further assumed that this current involves charged ions which undergo multiple collisions with air. These collisions transfer momentum. All of the measured data was consistent with this model. Many configurations were tested, and the results suggest general design principles for ACTs to be used for a variety of purposes.

Canning, Francis X.↗

NASA NTRS1973

Performance optimized, small structurally integrated ion thruster system

A 5-cm structurally integrated ion thruster has been developed for attitude control and stationkeeping of synchronous satellites. As optimized with a conventional ion extraction system, the system demonstrates a thrust T = 0.47 mlb at a beam voltage of 1600 V, total mass efficiency of 76%, and electrical efficiency of 56%. Under the subject contract effort, no significant performance change was noted for operation with two dimensional electrostatic thrust-vectoring grids. Structural integrity with the vectoring grids was demonstrated for shock (+ or - 30 G), sinusoidal (9 G), and random (19.9 G rms) accelerations. System envelope is 31.2 cm long by 13.4 cm flange bolt circle, with a mass of 9.0 Kg, including 6.8 Kg mercury propellant.

Hyman, J., Jr.↗

NASA NTRS2002

Momentum and Heat Flux Measurements in the Exhaust of VASIMR Using Helium Propellant

Electromagnetic thrusters typically use electric and magnetic fields to accelerate and exhaust plasma through interactions with the charged particles in the plasma. The energy required to create the plasma, i.e. ionization energy, is potential energy between the electron and ion. This potential energy is typically lost since it is not recovered as the plasma is exhausted and is known as frozen flow loss. If the frozen flow energy is a small fraction of the total plasma energy, this frozen flow loss may be negligible. However, if the frozen flow energy is a major fraction of the total plasma energy, this loss can severely reduce the energy efficiency of the thruster. Recovery and utilization of this frozen flow energy can improve the energy efficiency of a thruster during low specific impulse operating regimes when the ionization energy is a large fraction of the total plasma energy. This paper quantifies the recovery of the frozen flow energy, i.e. recombination energy, via the process of surface recombination for helium. To accomplish this task the momentum flux and heat flux of the plasma flow were measured and compared to calculated values from electrostatic probe data. This information was used to deduce the contribution of recombination energy to the total heat flux on a flat plate as well as to characterize the plasma conditions. Helium propellant was investigated initially due to its high ionization potential and hence available recombination energy.

Chavers, D. Gregory↗

NASA NTRS2007

The New NASA-STD-4005 and NASA-HDBK-4006, Essentials for Direct-Drive Solar Electric Propulsion

High voltage solar arrays are necessary for direct-drive solar electric propulsion, which has many advantages, including simplicity and high efficiency. Even when direct-drive is not used, the use of high voltage solar arrays leads to power transmission and conversion efficiencies in electric propulsion Power Management and Distribution. Nevertheless, high voltage solar arrays may lead to temporary power disruptions, through the so-called primary electrostatic discharges, and may permanently damage arrays, through the so-called permanent sustained discharges between array strings. Design guidance is needed to prevent these solar array discharges, and to prevent high power drains through coupling between the electric propulsion devices and the high voltage solar arrays. While most electric propulsion systems may operate outside of Low Earth Orbit, the plasmas produced by their thrusters may interact with the high voltage solar arrays in many ways similarly to Low Earth Orbit plasmas. A brief description of previous experiences with high voltage electric propulsion systems will be given in this paper. There are two new official NASA documents available free through the NASA Standards website to help in designing and testing high voltage solar arrays for electric propulsion. They are NASA-STD-4005, the Low Earth Orbit Spacecraft Charging Design Standard, and NASA-HDBK-4006, the Low Earth Orbit Spacecraft Charging Design Handbook. Taken together, they can both educate the high voltage array designer in the engineering and science of spacecraft charging in the presence of dense plasmas and provide techniques for designing and testing high voltage solar arrays to prevent electrical discharges and power drains.

Ferguson, Dale C.↗

NASA NTRS2007

Space Technology 7 : Micropropulsion and Mass Distribution

The NASA New Millennium Program Space Technology 7 (ST7) project will validate technology for precision spacecraft control. The ST7 disturbance reduction system (DRS) will contain new micropropulsion technology to be flown as part of the European Space Agency's LISA (laser interferometer space antenna) Pathfinder project. After launch into a low Earth orbit in early 2010, the LISA Pathfinder spacecraft will be maneuvered to a halo orbit about the Earth-Sun LI Lagrange point for operations. The DRS will control the position of the spacecraft relative to a reference to an accuracy of one nanometer over time scales of several thousand seconds. To perform the control the spacecraft will use a new colloid thruster technology. The thrusters will operate over the range of 5 to 30 micro-Newtons with precision of 0.1 micro-Newton. The thrust will be generated by using a high electric field to extract charged droplets of a conducting colloid fluid and accelerating them with a precisely adjustable voltage. The control position reference will be provided by the European LISA Technology Package, which will include two nearly free-floating test masses. The test mass position and attitude will be sensed and adjusted using electrostatic capacitance bridges. The DRS will control the spacecraft position with respect to one test mass while minimizing disturbances on the second test mass. The dynamic control system will cover eighteen degrees of freedom, six for each of the test masses and six for the spacecraft. In the absence of other disturbances, the test masses will slowly gravitate toward local concentrations of spacecraft mass. The test mass acceleration must be minimized to maintain the acceleration of the enclosing drag-free spacecraft within the control authority of the micropropulsion system. Therefore, test mass acceleration must be predicted by accurate measurement of mass distribution, then offset by the placement of specially shaped balance masses near each test mass. The - acceleration is characterized by calculating the gravitational effect of over ten million modeled points of a nearly 500-kg spacecraft. This paper provides an overview of the mission technology and the process of precision mass modeling of the DRS equipment.

ST-7↗

NASA NTRS1993

Inertial electrostatic confinement as a power source for electric propulsion

The potential use of an INERTIAL ELECTROSTATIC CONFINEMENT (IEC) power source for space propulsion has previously been suggested by the authors and others. In the past, these discussions have generally followed the charged-particle electric-discharge engine (QED) concept proposed by Bussard, in which the IEC is used to generate an electron beam which vaporizes liquid hydrogen for use as a propellant. However, an alternate approach is considered, using the IEC to drive a 'conventional' electric thruster unit. This has the advantage of building on the rapidly developing technology for such thrusters, which operate at higher specific impulse. Key issues related to this approach include the continued successful development of the physics and engineering of the IEC unit, as well as the development of efficient step-down dc voltage transformers. The IEC operates by radial injection of energetic ions into a spherical vessel. A very high ion density is created in a small core region at the center of the vessel, resulting in extremely high fusion power density in the core. Experiments at the U. of Illinois in small IEC devices (is less than 60 cm. dia.) demonstrated much of the basic physics underlying this concept, e.g. producing 10(exp 6) D-D neutrons/sec steady-state with deuterium gas flow injection. The ultimate goal is to increase the power densities by several orders of magnitude and to convert to D-He-3 injection. If successful, such an experiment would represent a milestone proof-of-principle device for eventual space power use. Further discussion of IEC physics and status are presented with a description of the overall propulsion system and estimated performance.

Miley, G. H.↗

NASA NTRS1991

Characteristics of ring-cusp discharge chambers

Measurements have been obtained for the operating characteristics of a 30 cm diameter ring-cusp ion thruster (RCIT), quantitatively comparing its performance parameters to those of a divergent-field J-series cluster of the same size. The high level of performance established for the RCIT is due to its maintenance of both a higher primary-electron population and Maxwellian-electron temperature, as the beam-ion production cost is reduced to its baseline value. Ion losses to the discharge-chamber walls can be reduced by an applied electrostatic field.

Matossian, J. N.↗

NASA NTRS1993

Inertial electrostatic confinement as a power source for electric propulsion

The potential use of an Inertial Electrostatic Confinement (IEC) power source for space propulsion has previously been suggested by the authors and others. In the past, these discussions have generally followed the charged-particle electric-discharge engine (QED) concept proposed by Bussard, in which the IEC is used to generate an electron beam which vaporizes liquid hydrogen for use as a propellant. However, in the present study, we consider an alternate approach, using the IEC to drive a conventional electric thruster unit. This has the advantage of building on the rapidly developing technology for such thrusters, which operate at higher specific impulse. Key issues related to this approach include the continued successful development of the physics and engineering of the IEC unit, as well as the development of efficient step-down dc voltage transformers. The IEC operates by radial injection of energetic ions into a spherical vessel. A very high ion density is created in a small core region at the center of the vessel, resulting in extremely high fusion power density in the core. Present experiments at the U. of Illinois in small IEC devices (less than 60-cm. dia.) have demonstrated much of the basic physics underlying this concept, e.g. producing approximately 10(exp 6) D-D neutrons/sec steady-state with deuterium gas flow injection. The ultimate goal is to increase the power densities by several orders of magnitude and to convert to D-He-3 injection. If successful, such an experiment would represent a milestone proof-of-principle device for eventual space power use. Further discussion of IEC physics and status will be presented with a description of the overall propulsion system and estimated performance.

Miley, George H.↗