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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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All sourcesNASA NTRSDOE OSTI
NASA NTRS1986

Ground testing in a simulated auroral environment

Shuttle polar orbit missions are being planned in which astronaut extravehicular activities (EVA) will eventually include external operations or repair and refurbishment of satellites. Recent satellite data and model calculations show that significant spacecraft charging and differential charging occurs in that environment and could endanger equipment and personnel involved in the EVA. These spacecraft charging events take place in response to an intense flux of high energy electrons accompanied by a large drop in the ambient plasma density. The extent of which the auroral environment can be simulated in a large vacuum chamber are determined so that conditions which lead to significant charging and discharging may be observed and controlled. A large chamber would be required in order to simulate the effects of an aurora-like, large cross-section, energetic electron precipitation on a target system the size of an astronaut with a life support system. The precipitating electrons can be simulated with an array of multipactor sources and the background ionization produced by the beam can be limited to the range of the auroral density if the chamber pressure is not significantly above 1 x .000001 Torr. Some small chamber work has shown that the plasma in the beam will diffuse radially at a rate determined by the Bohm diffusion coefficient.

Mcintyre, Bernard↗
NASA NTRS1988

An unusual charging event on ISEE 1

Electrostatic cleanliness requirements on ISEE 1 were expected to prevent negative charging in sunlight. This has largely been true, but on three occasions, ISEE 1 has been observed to charge to significant negative potentials in sunlight. Data from the two electric field experiments and from the plasma composition experiment on ISEE 1 show that the spacecraft charged to close to -70 V in sunlight at about 0700 UT on March 17, 1978. Data from the electron spectrometer experiment show that there was a potential barrier of some -10 to -20 V about the spacecraft during this event. The potential barrier was effective in turning back emitted photoelectrons to the spacecraft. Potential barriers can be formed by differential charging on the spacecraft or by the presence of excess space charge in the plasma. The shape of the barrier suggests that it is due to the former, even though electrostatic cleanliness specifications imposed on ISEE were intended to eliminate differential charging. Modeling of this event showed that the barrier could not be produced by the presence of space charge but that it was most likely produced by differential charging of the solar arrays.

Olsen, R. C.↗
DOE OSTI2025

Measurements of differential charged-current cross sections on argon for electron neutrinos with final-state protons in MicroBooNE

This work presents single-differential electron-neutrino charged-current cross sections on argon measured using the MicroBooNE detector at the Fermi National Accelerator Laboratory. The analysis uses data recorded when the Neutrinos at the Main Injector beam was operating in both neutrino and antineutrino modes, with exposures of $2 \times 10^{20}$ and $5 \times 10^{20}$ protons on target, respectively. A selection algorithm targeting electron-neutrino charged-current interactions with at least one proton, one electron, and no pions in the final topology is used to measure differential cross sections as a function of outgoing electron energy, total visible energy, and opening angle between the electron and the most energetic proton. The interaction rate as a function of proton multiplicity is also reported. The total cross section is measured as [4.1 $\pm$ 0.4 (stat.) $\pm$ 1.2 (syst.)]$ $$\times 10^{-39} \mathrm{cm}^{2}/ \mathrm{nucleon}$. The unfolded cross-section measurements are compared to predictions from neutrino event generators commonly employed in the field. Good agreement is seen across all variables within uncertainties.

Abratenko, P. [Tufts U. (main)]↗
NASA NTRS1977

Conductive spacecraft materials development program

The objectives of this program are to provide design criteria, techniques, materials, and test methods to ensure control of absolute and differential charging of spacecraft surfaces. The control of absolute and differential charging of spacecraft cannot be effected without the development of new and improved or modified materials or techniques that will provide electrical continuity over the surface of the spacecraft. The materials' photoemission, secondary emission, thermooptical, physical, and electrical properties in the space vacuum environment both in the presence and absence of electrical stress and ultraviolet, electron, and particulate radiation, are important to the achievement of charge control. The materials must be stable or have predictable response to exposure to the space environment for long periods of time. The materials of interest include conductive polymers, paints, transparent films and coatings as well as fabric coating interweaves.

Lehn, W. L.↗
NASA NTRS2015

On Possible Arc Inception on Low Voltage Solar Array

Recent analysis of spacecraft failures during the period of 1990-2013 demonstrated clearly that electrostatic discharges caused more than 8 of all registered failures and anomalies, and comprised the most costly losses (25) for operating companies and agencies. The electrostatic discharges on spacecraft surfaces are the results of differential charging above some critical (threshold) voltages. The mechanisms of differential charging are well known, and various methods have been developed to prevent a generation of significant electric fields in areas of triple junctions. For example, low bus voltages in Low Earth Orbit plasma environment and slightly conducting layer over coverglass (ITO) in Geosynchronous Orbit surroundings are believed to be quite reliable measures to prevent discharges on respective surfaces. In most cases, the vulnerable elements of spacecraft (solar arrays, diode boards, etc.) go through comprehensive ground tests in vacuum chambers. However, tests articles contain the miniscule fragments of spacecraft components such as 10-30 solar cells of many thousands deployed on spacecraft in orbit. This is one reason why manufacturing defects may not be revealed in ground tests but expose themselves in arcing on array surface in space. The other reason for ineffectiveness of discharge preventive measures is aging of all materials in harsh orbital environments. The expected life time of modern spacecraft varies within the range of five-fifteen years, and thermal cycling, radiation damages, and mechanical stresses can result in surface erosion on conductive layers and microscopic cracks in coverglass sheets and adhesive films. These possible damages may cause significant increases in local electric field strengths and subsequent discharges. The primary discharges may or may not be detrimental to spacecraft operation, but they can produce the necessary conditions for sustained arcs initiation. Multiple measures were developed to prevent sustained discharges between adjacent strings, and many ground tests were performed to determine threshold parameters (voltage and current) for sustained arcs. And again, manufacturing defects and aging in space environments may result in considerable decrease of critical threshold parameters. This paper is devoted to the analysis of possible reasons behind arcing on spacecraft with low bus voltages.

plasma interaction↗
NASA NTRS2015

On Possible Arc Inception on Low Voltage Solar Array

Recent analysis of spacecraft failures during the period of 1990-2013 demonstrated clearly that electrostatic discharges caused more than 8 percent of all registered failures and anomalies, and comprised the most costly losses (25 percent) for operating companies and agencies. The electrostatic discharges on spacecraft surfaces are the results of differential charging above some critical (threshold) voltages. The mechanisms of differential charging are well known, and various methods have been developed to prevent a generation of significant electric fields in areas of triple junctions. For example, low bus voltages in Low Earth Orbit plasma environment and slightly conducting layer over cover-glass (ITO) in Geosynchronous Orbit surroundings are believed to be quite reliable measures to prevent discharges on respective surfaces. In most cases, the vulnerable elements of spacecraft (solar arrays, diode boards, etc.) go through comprehensive ground tests in vacuum chambers. However, tests articles contain the miniscule fragments of spacecraft components such as 10-30 solar cells of many thousands deployed on spacecraft in orbit. This is one reason why manufacturing defects may not be revealed in ground tests but expose themselves in arcing on array surface in space. The other reason for ineffectiveness of discharge preventive measures is aging of all materials in harsh orbital environments. The expected life time of modern spacecraft varies within the range of five-fifteen years, and thermal cycling, radiation damages, and mechanical stresses can result in surface erosion on conductive layers and microscopic cracks in cover-glass sheets and adhesive films. These possible damages may cause significant increases in local electric field strengths and subsequent discharges. The primary discharges may or may not be detrimental to spacecraft operation, but they can produce the necessary conditions for sustained arcs initiation. Multiple measures were developed to prevent sustained discharges between adjacent strings, and many ground tests were performed to determine threshold parameters (voltage and current) for sustained arcs. And again, manufacturing defects and aging in space environments may result in considerable decrease of critical threshold parameters. This paper is devoted to the analysis of possible reasons behind arcing on spacecraft with low bus voltages.

low voltage solar array↗
NASA NTRS2016

Solar Array at Very High Temperatures: Ground Tests

Solar array design for any spacecraft is determined by the orbit parameters. For example, operational voltage for spacecraft in Low Earth Orbit (LEO) is limited by significant differential charging due to interactions with low temperature plasma. In order to avoid arcing in LEO, solar array is designed to generate electrical power at comparatively low voltages (below 100 V) or to operate at higher voltages with encapsulated of all suspected discharge locations. In Geosynchronous Orbit (GEO) differential charging is caused by energetic electrons that produce differential potential between coverglass and conductive spacecraft body in a kilovolt range. In such a case, weakly conductive layer over coverglass (ITO) is one of possible measures to eliminate dangerous discharges on array surface. Temperature variations for solar arrays in both orbits are measured and documented within the range of -150 C +110 C. This wide interval of operational temperatures is regularly reproduced in ground tests with radiative heating and cooling inside shroud with flowing liquid nitrogen. The requirements to solar array design and tests turn out to be more complicated when planned trajectory crosses these two orbits and goes closer to Sun. Conductive layer over coverglass causes sharp increase in parasitic current collected from LEO plasma, high temperature may cause cracks in encapsulating material (RTV), radiative heating of coupon in vacuum chamber becomes practically impossible above 150 C, conductivities of glass and adhesive go up with temperature that decrease array efficiency, and mechanical stresses grow up to critical magnitudes. A few test arrangements and respective results are presented in current paper. Coupons were tested against arcing in simulated LEO and GEO environments under elevated temperatures up to 200 C. The dependence of leakage current on temperature was measured, and electrostatic cleanness was verified for coupons with antireflection (AR) coating over ITO layer.

differential charging↗
NASA NTRS2016

Solar Array at Very High Temperatures: Ground Tests

Solar array design for any spacecraft is determined by the orbit parameters. For example, operational voltage for spacecraft in Low Earth Orbit (LEO) is limited by significant differential charging due to interactions with low temperature plasma. In order to avoid arcing in LEO, solar array is designed to generate electrical power at comparatively low voltages (below 100 volts) or to operate at higher voltages with encapsulation of all suspected discharge locations. In Geosynchronous Orbit (GEO) differential charging is caused by energetic electrons that produce differential potential between the coverglass and the conductive spacecraft body in a kilovolt range. In such a case, the weakly conductive layer over coverglass, indium tin oxide (ITO) is one of the possible measures to eliminate dangerous discharges on array surface. Temperature variations for solar arrays in both orbits are measured and documented within the range of minus150 degrees Centigrade to plus 1100 degrees Centigrade. This wide interval of operational temperatures is regularly reproduced in ground tests with radiative heating and cooling inside a shroud with flowing liquid nitrogen. The requirements to solar array design and tests turn out to be more complicated when planned trajectory crosses these two orbits and goes closer to the Sun. The conductive layer over coverglass causes a sharp increase in parasitic current collected from LEO plasma, high temperature may cause cracks in encapsulating (Room Temperature Vulcanizing (RTV) material; radiative heating of a coupon in vacuum chamber becomes practically impossible above 1500 degrees Centigrade; conductivities of glass and adhesive go up with temperature that decrease array efficiency; and mechanical stresses grow up to critical magnitudes. A few test arrangements and respective results are presented in current paper. Coupons were tested against arcing in simulated LEO and GEO environments under elevated temperatures up to 2000 degrees Centigrade. The dependence of leakage current on temperature was measured, and electrostatic cleanness was verified for coupons with antireflection (AR) coating over the indium tin oxide (ITO) layer.

differential charging↗
NASA NTRS2018

Modeling of Orbital and Attitude Dynamics of Nanosatellites Controlled via Active Electrostatic Charging

The large-scale exploration of airless bodies, such as asteroids and moons, is gaining interest, however it is limited by mobility issues: the lack of atmosphere, low gravity, and unknown soil properties pose difficult challenges for many forms of traditional locomotion. The environment in proximity of these bodies is also electrically charged due to interactions with solar wind and UV radiation. The EGlider (Electrostatic Glider) concept would be able to overcome these mobility issues by leveraging the natural environment, allowing operations in close proximity of the surface, while enabling long duration missions by minimizing propellant consumption. The EGlider is an advanced concept for small satellite mobility and propulsion, which relies on the electric fields naturally present around airless bodies in order to generate forces and torques useful for maneuvering. It does so by extending electrically charged appendages, which enable it to electrostatically soar above the surface. By differentially charging its electrodes it can also produce torques to control its attitude. The charges are maintained by continuous active ion or electron emission from the spacecraft, in order to cancel out the neutralizing influx of charges from surrounding plasma. An investigation of the spacecraft-plasma interaction was carried out. This included studying the effect of electrode geometry and calculating the charge-to-mass ratios required to enable several mission scenarios. Long, thin wire electrodes were identified to be the most power-efficient and would allow power-to-weight ratios achievable with current nanosatellite technologies. High electrode potential represents the main limiting factor for the system design. In order to test the feasibility of active control by means of differential charging, a simple 2D interaction model was developed, and a feedback controller to stabilize the vehicle was tested in a simulation environment. The results confirmed that good performance could be obtained for both position and attitude control. Finally, a dedicated software was developed for future simulation and testing of control strategies for the EGlider. This software allows to study the trajectory and attitude of an arbitrarily configured spacecraft in the proximity of an arbitrarily defined main airless body. The spacecraft can be assembled from basic parts, each with specified electrical, mass and optical properties. Efficient models allow to calculate gravitational and electrical interactions with the rotating main body and the local plasma field, as well as solar radiation pressure effects. Control models can be implemented as simple plug-in functions and easily tested. The preliminary validation campaign showed good matching with the reference cases that have been analyzed.

Corpino, Sabrina↗
DOE OSTI2025

Measurement of double-differential charged-current Drell-Yan cross-sections at high transverse masses in $pp$ collisions at $\sqrt{s}$ = 13 TeV with the ATLAS detector

This paper presents a first measurement of the cross-section for the charged-current Drell-Yan process pp → W ± → ℓ ± ν above the resonance region, where ℓ is an electron or muon. The measurement is performed for transverse masses, $m$$^{W}_{T}$, between 200 GeV and 5000 GeV, using a sample of 140 fb −1 of pp collision data at a centre-of-mass energy of = 13 TeV collected by the ATLAS detector at the LHC during 2015–2018. The data are presented single differentially in transverse mass and double differentially in transverse mass and absolute lepton pseudorapidity. A test of lepton flavour universality shows no significant deviations from the Standard Model. The electron and muon channel measurements are combined to achieve a total experimental precision of 3% at low $m$$^{W}_{T}$. The single- and double differential W-boson charge asymmetries are evaluated from the measurements. A comparison to next-to-next-to-leading-order perturbative QCD predictions using several recent parton distribution functions and including next-to-leading-order electroweak effects indicates the potential of the data to constrain parton distribution functions. The data are also used to constrain four fermion operators in the Standard Model Effective Field Theory formalism, in particular the lepton-quark operator Wilson coefficient $c$$^{(c)}_{ℓq}$.

Hadron-Hadron Scattering↗
NASA NTRS1984

Anomalously high potentials observed on ISEE

Data from the two electric field experiments and from the plasma composition experiment on ISEE-1 show that the spacecraft charged to close to -70 V in sunlight at about 0700 UT on March 17, 1978. Data from the electron spectrometer experiment show that there was a potential barrier of some -10 to -20 V about the spacecraft during this event. The potential barrier was effective in turning back emitted photoelectrons to the spacecraft. Potential barriers can be formed because of differential charging on the spacecraft or because of the presence of space charge. The stringent electrostatic cleanliness specifications imposed on ISEE make the presence of differential charging unlikely, if these precautions were effective. Modeling of this event is required to determine if the barrier was produced by the presence of space charge.

Whipple, E. C.↗
NASA NTRS1977

Charge-coupled differential amplifier

Device solves problem of differential voltage gain in sampled-data metal-oxide semiconductor, producing gain at any frequency compatible with CCD operation, with low noise, low power, good linearity, and high stability.

Hewes, C. R.↗
NASA NTRS1978

Differential photoelectric charging of nonconducting surfaces in space

The photoelectric charging caused by an infinitely long strip of sunlight across a nonconducting plane is studied by use of a model which contains an electrical cutoff radius, and the results of numerical calculations are presented. The model simulates charging of a sunlit area with dimensions equal to the strip's width, exposed to a plasma with a comparatively large Debye length. Uniform potential is quickly established on a uniformly sunlit strip as a result of charge redistribution by low-energy photoelectrons. The results are in accord with a theoretical surface conductivity derived for photoelectron sheaths above highly charged sunlit areas. The surface potential, which drops sharply across the sunlight-shadow boundary, is discussed.

Pelizzari, M. A.↗
NASA NTRS1980

Configuration effects on satellite charging response

The response of various spacecraft configurations to a charging environment in sunlight was studied using the NASA Charging Analyzer Program code. The configuration features geometry, type of stabilization, and overall size. Results indicate that sunlight charging response is dominated by differential charging effects. Shaded insulation charges negatively result in the formation of potential barriers which suppress photoelectron emission from sunlit surfaces. Sunlight charging occurs relatively slowly: with 30 minutes of charging simulations, in none of the configurations modeled did the most negative surface cell reach half its equilibrium potential in eclipse.

Purvis, C. K.↗
NASA NTRS1985

Electric propulsion produced environments and possible interactions with the SP-100 power system

An examination of a conceptual SP-100 type power system has been conducted to assess the effect of possible electric thruster induced contamination. The operating characteristics and known exhaust plume properties of various electric thrusters are summarized along with the natural environment and its effects to provide background for the contamination discussion. Charge buildup on nonconducting surfaces, differential charging of surfaces and arcing should be moderated by the operation of an electric thruster system. Thin film deposition of eroded thruster materials is a possible contamination mechanism. Electric thruster field efflux and propellant condensation should have only a minor contaminating effect upon a nuclear power system.

Deininger, W. D.↗
NASA NTRS2023

Lessons Learned From Evaluation and Mitigation of Space Charging Threat Due to Use of Isolated (Hybrid) Bearings on the Pace Ocean Color Instrument

A high precision, high resolution Ocean Color Instrument (OCI) was developed for the Plankton, Aerosol, Cloud ocean Ecosystem (PACE) mission which required the use of electrically isolated Hybrid Bearings (silicon nitride balls). While desirable for technical performance of the instrument, this application caused portions of the instrument to become electrically isolated and come under threat of space charging effects from charged particles and the resultant differential charging. An evaluation of the environment, susceptibility of the components, such as bearings and electronics, and mitigation strategies were performed. As part of that process, simulations of the charging environment, analysis of internal mechanism charging and prediction of most likely discharge paths, testing of the bearings for susceptibility as well as resulting damage were performed and analyzed. Additionally, a novel method to deal with charging was developed. Considerations for dealing with electrically isolated mechanisms, results of the above-mentioned efforts as well as lessons learned are presented.

space charging↗
NASA NTRS1981

The Auroral Particles experiment

An instrument for the detection of particles in the energy range of 0.1 ev to 80 Kev was designed, built, tested, calibrated, and flown onboard the spacecraft ATS-6. Data from this instrument generated the following research: intensive studies of the plasma in the vicinity of the spacecraft; global variations of plasmas; correlative studies using either other spacecraft or ground based measurements; and studies of spacecraft interactions with ambient plasmas including charging, local electric fields due to differential charging, and active control of spacecraft potential. Results from this research are presented.

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