SEARCH · Search NASA
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.
Quote a phrase for an exact phrase match. Source license links do not imply unrestricted reuse.
Attempt was made to determine solar cell panel peak power capability without disrupting power flow from panel. Separate solar cell strings were switched from panel circuits, and increasingly larger loads were added rapidly until peak power points were transversed. String wattage output was recorded and all stored string measurements summed together indicate peak power point in panel.
The technological readiness of the proposed process sequence was reviewed. Process steps evaluated include: (1) plasma etching to establish a standard surface; (2) forming junctions by diffusion from an N-type polymeric spray-on source; (3) forming a p+ back contact by firing a screen printed aluminum paste; (4) forming screen printed front contacts after cleaning the back aluminum and removing the diffusion oxide; (5) cleaning the junction by a laser scribe operation; (6) forming an antireflection coating by baking a polymeric spray-on film; (7) ultrasonically tin padding the cells; and (8) assembling cell strings into solar circuits using ethylene vinyl acetate as an encapsulant and laminating medium.
The use of high-voltage solar arrays in space is discussed in connection with the draining of array power by currents flowing between exposed surfaces through the surrounding plasma. The possibility of reducing the power loss by arranging solar cell strings in repeated small-area modules to eliminate any large areas at high potentials is investigated. It is found that the difference in power loss between modular and linear patterned high-voltage arrays is fairly small. Although the use of modular patterns can reduce the effective mean potential by about 10%, for the type of configuration being considered there is also a 10% increase in sheath area, leading to only a few percent change in total power loss. It is concluded that plasma power loss should not be a primary consideration in designing the physical arrangement of high-voltage arrays.
Artificial strings of solar oscillation data with gaps and noise, corresponding to the output of different spatial filter functions, were analyzed. Peaks in the power spectrum are identified for values of the degree l from 0 to 18, and rotational splitting is estimated. The filters prove effective in facilitating identification of essentially all the real peaks in the power spectrum. Estimates of peak frequencies and amplitudes and rotational splitting frequencies are in reasonably good agreement with the input values. Spurious peaks in autocorrelation spectra correspond to the frequency spacing between power peaks with the same order n, differing by one or two in the degree l.
The main objective was to design and build a minimum of three photovoltaic test panels for plasma interaction experiments. These experiments are intended to provide data on the interactions between high-voltage solar arrays and the space plasma environment. Data gathered will significantly contribute to the development of design criteria for the space station solar arrays. Electrical isolation between the solar cell strings and the module mounting plate is required for high-voltage bias.
Space environmental effects have caused severe problems as satellites move toward increased power and operating voltage levels. The greatest unknown, however, is the effect of high velocity micrometeoroid impacts on high voltage arrays (>200V). Understanding such impact phenomena is necessary for the design of future reliable, high voltage solar arrays, especially for Space Solar Power applications. Therefore, the objective of this work was to study the effect of hypervelocity impacts on high voltage solar arrays. Initially, state of the art, 18% efficient GaAs solar cell strings were targeted. The maximum bias voltage on a two-cell string was -200 V while the adjacent string was held at -140 V relative to the plasma potential. A hollow cathode device provided the plasma. Soda lime glass particles 40-120 micrometers in diameter were accelerated in the Hypervelocity Impact Facility to velocities as high as 11.6 km/sec. Coordinates and velocity were obtained for each of the approximately 40 particle impact sites on each shot. Arcing did occur, and both discharging and recharging of arcs between the two strings was observed. The recharging phenomena appeared to stop at approximately 66V string differential. No arcing was observed at 400 V on concentrator cell modules for the Stretched Lens Array.
A spacecraft in a high-density equatorial LEO plasma will float negative relative to the ambient plasma. Because of the electron collection of exposed conductors on its solar arrays, it may float negative by up to its array voltage. The floating potential depends on the relative areas of electron and ion collection of the spacecraft. Early estimates of the International Space Station (ISS) potential were about -140 V relative to the surrounding plasma, because of its 160 V solar array string voltage. Because of the possibility of arcing of ISS structures and astronaut EMUs (spacesuits) into the space plasma, Plasma Contacting Units (PCUs) were added to the ISS design, to reduce the highly negative floating potentials by emitting electrons (effectively increasing the ion collecting area). In addition to the now-operating ISS PCUs, safety rules require another independent arc-hazard control method. In this paper, I discuss alternatives to the ISS PCUs for keeping the ISS floating potential at values below the arc-thresholds of ISS and EMU surface materials. Advantages and disadvantages of all of the recline loss will be presented.
During the last few years of Solar Minimum, GEO spacecraft charging design practices may have become lax because of paucity of spacecraft charging events. Unfortunately, this has also been the time of great changes in spacecraft design, because of the new emphases on higher power arrays and lower costs. Also unfortunate is the fact that spacecraft charging may lead to failures of solar array strings, panels, or entire spacecraft. One way to prevent satellite failures die to spacecraft charging events is to simulate the effects with a charging code, such as the venerable NASCAP/GEO code. We will discuss the use of NASCAP on the ACTS satellite as well as a newer application dealing with typical recent spacecraft charging anomalies.
This article presents a tri-port current-source soft-switching medium-voltage string inverter (TMVSI) to reduce the Levelized cost of energy (LCOE) of large-scale solar-plus-storage (SPS) farms. Throughout this work, the topology, operating principle, circuit simulation, control, and hardware design of the TMVSI are introduced. In addition, three key challenges compromising the performance of the TMVSI are addressed in this article. First, a new feed-forward compensation for model predictive control of the TMVSI is proposed to compensate for sampling and computational delay and high dc-link ripple with low computational cost and high scalability. Second, the use of laminated permanent magnets is proposed to reduce the losses of the MV medium-frequency transformer (MFT) in the TMVSI while increasing its saturation current. Finally, an improved U-shape winding pattern is proposed to decrease the leakage inductance of the MV MFT and reduce the voltage stress across semiconductor switches without adding cost or complexity. Here, the effectiveness of the TMVSI is validated by experiments at up to 20kW/500V in different test cases with custom-built prototypes. The contributions of this work make the TMVSI a viable solution to SPS farms and unleash its potential for LCOE reduction.
It has been known for many years, based partly on NASA Glenn Research Center testing, that high-voltage solar arrays arc into the space plasma environment. Solar arrays are composed of solar cells in series with each other (a string), and the strings may be connected in parallel to produce the entire solar array power. Arcs on solar arrays can damage or destroy solar cells, and in the extreme case of sustained arcing, entire solar array strings, in a flash. In the case of sustained arcing (discovered at Glenn and applied to the design and construction of solar arrays on Space Systems/Loral (SS/Loral, Palo Alto, CA) satellites, Deep-Space 1, and Terra), an arc on one solar array string can couple to an adjacent string and continue to be powered by the solar array output until a permanent electrical short is produced. In other words, sustained arcs produced by arcs into the plasma (so-called trigger arcs) may turn into disastrous sustained arcs by involving other array strings.
Spacecraft with solar arrays operate throughout the near earth environment and are increasingly planned for outer planet missions. An often overlooked test condition for solar arrays that is applicable to these missions is micrometeorite impacts and possibly electrostatic discharge (ESD) events resulting from these impacts. The Marshall Space Flight Center (MSFC) is partnering with Space Systems/Loral, LLC (SSL) to examine the results of simulated micrometeorite impacts on the electrical performance of an advanced, lightweight flexible solar array design. The test is performed at NASA MSFC's Microlight Gas Gun Facility. The SSL-provided coupons consist of three strings, each string with two solar cells in series. Five impacts will be induced at various locations on a powered test coupon under different string voltage (0 volts - 150 volts) and string current (1.1 amperes - 1.65 amperes) conditions. The maximum specified test voltage and current represent margins of 1.5 times for both voltage and current. The test parameters are chosen to demonstrate new array design robustness to any ESD event caused by plasma plumes resulting from a simulated micrometeorite impact. A second unpowered coupon will undergo two impacts: one impact on the front side and one impact on the back side. Following the impact testing, the second coupon will be exposed to a thermal cycle test to determine possible damage propagation and further electrical degradation due to thermally-induced stress. The setup, checkout, and results from the impact testing are discussed. The challenges for impact testing include precise coupon alignment to control impact location; pressure management during the impact process; and measurement of the true transient electrical response during impact on the powered coupon. Results from pre- and post-test visual and electrical functional testing are also discussed.
Photovoltaic (PV) modules consist of a string of electrically interconnected silicon solar cells capable of producing practical quantities of electrical power when exposed to sunlight. To insure high reliability and long term performance, the functional components of the solar cell module must be adequately protected from the environment by some encapsulation technique. The encapsulation system must provide mechanical support for the cells and corrosion protection for the electrical components. The goal of the program is to identify and develop encapsulation systems consistent with the PV module operating requirements of 30 year life and a target cost of $0.70 per peak watt ($70 per square meter) (1980 dollars). Assuming a module efficiency of ten percent, which is equivalent to a power output of 100 watts per square meter in midday sunlight, the capital cost of the modules may be calculated to be $70.00 per square meter. Out of this cost goal, only 20 percent is available for encapsulation due to the high cost of the cells, interconnects, and other related components. The encapsulation cost allocation may then be stated as $14.00 per square meter, included all coatings, pottant and mechanical supports for the cells.
The successes of the Cassini/Huygens mission have heightened interest to return to the Saturn system with focused robotic missions. The desire for a sustained presence at Titan, through a dedicated orbiter and in-situ vehicle, either a lander or aerobot, has resulted in definition of a Titan Explorer flagship mission as a high priority in the Solar System Exploration Roadmap. The discovery of active water vapor plumes erupting from the tiger stripes on the moon Enceladus has drawn the attention of the space science community. The NASA's Evolutionary Xenon Thruster (NEXT) ion propulsion system is well suited to future missions to the Saturn system. NEXT is used within the inner solar system, in combination with a Venus or Earth gravity assist, to establish a fast transfer to the Saturn system. The NEXT system elements are accommodated in a separable Solar Electric Propulsion (SEP) module, or are integrated into the main spacecraft bus, depending on the mission architecture and performance requirements. This paper defines a range of NEXT system configurations, from two to four thrusters, and the Saturn system performance capability provided. Delivered mass is assessed parametrically over total trip time to Saturn. Launch vehicle options, gravity assist options, and input power level are addressed to determine performance sensitivities. A simple two-thruster NEXT system, launched on an Atlas 551, can deliver a spacecraft mass of over 2400 kg on a transfer to Saturn. Similarly, a four-thruster system, launched on a Delta 4050 Heavy, delivers more than 4000 kg spacecraft mass. A SEP module conceptual design, for a two thruster string, 17 kW solar array, configuration is characterized.
String theory in AdS 3 with purely NS-NS fluxes and vanishing RR moduli has a continuum of winding string excitations in radial plane wave states. BTZ black holes can emit such strings, which then flow out toward the AdS 3 boundary as a stream of massive quanta, and form a black hole analogue of the solar wind. The winding string sector thus provides a decay channel for the black hole to evaporate without having either to couple the system to an external reservoir or to match the AdS 3 throat onto an asymptotically flat region. We compute the emission amplitude of this “holar wind” in the semi-classical approximation, and consider the associated version of the black hole information paradox.
This paper documents testing and analyses to quantify International Space Station (ISS) Solar Array Wing (SAW) string electrical performance under highly off-nominal, low-temperature-low-intensity (LILT) operating conditions with nonsolar light sources. This work is relevant for assessing feasibility and risks associated with a Sequential Shunt Unit (SSU) remove and replace (R&R) Extravehicular Activity (EVA). During eclipse, SAW strings can be energized by moonlight, EVA suit helmet lights or video camera lights. To quantify SAW performance under these off-nominal conditions, solar cell performance testing was performed using full moon, solar simulator and Video Camera Luminaire (VCL) light sources. Test conditions included 25 to 110 C temperatures and 1- to 0.0001-Sun illumination intensities. Electrical performance data and calculated eclipse lighting intensities were combined to predict SAW current-voltage output for comparison with electrical hazard thresholds. Worst case predictions show there is no connector pin molten metal hazard but crew shock hazard limits are exceeded due to VCL illumination. Assessment uncertainties and limitations are discussed along with operational solutions to mitigate SAW electrical hazards from VCL illumination. Results from a preliminary assessment of SAW arcing are also discussed. The authors recommend further analyses once SSU, R&R, and EVA procedures are better defined.
The rovers of the National Aeronautics and Space Administration’s (NASA) Mars Exploration Rovers (MER) project, Spirit (MER-A) and Opportunity (MER-B), safely landed on the surface of Mars three weeks apart during January 2004. Spirit and Opportunity were built and operated by the Jet Propulsion Laboratory (JPL), which is managed by the California Institute of Technology (Caltech) for NASA. Spirit landed at Gusev Crater, 14.8 degrees south of the equator, and operated continuously on the surface of Mars from January 4, 2004 until last contact from Spirit on March 22, 2010. Opportunity landed at Meridiani Planum, a location 2.5 degrees south of the Martian equator. Opportunity operated continuously on the Martian surface from January 25, 2004 until the last received transmission from Opportunity on June 10, 2018. The goal of the MER project was to determine if Mars ever had a habitable environment, in particular, if it ever had water. During their missions, both Spirit and Opportunity found evidence that liquid water once flowed on the surface of Mars. Both Spirit and Opportunity used a 1.33 m2 triple-junction solar array as their power sources. Based on observations of the original Mars rover, the Sojourner rover of the Mars Pathfinder mission that landed on Mars on July 4, 1997, the expectation was that the Martian dust would rapidly accumulate on the solar arrays of Spirit and Opportunity, and that the rovers would not have enough energy to continue operations after 90 Martian days (sols). Instead, due in part to lower dust accumulation rates than expected and numerous dust cleaning events, the Spirit and Opportunity rovers continued to operate on the Martian surface for over 2000 sols (MER-A) and 5000 sols (MER-B), respectively. During this time, the rovers experienced multiple Martian winters and several dust storms. Because the sources of solar array energy loss were known, the solar array energy output offered a method to scientifically estimate the loading and aeolian removal of dust from the solar arrays each sol. The MER Power operations team called this value that they calculated the solar array Dust Factor (DF). Dust Factor was defined as the fraction of sunlight that penetrates the accumulated dust on the surface of the solar array. A Dust Factor of 1.0 would indicate that the solar array was perfectly clean. A Dust Factor of 0.6 would indicate that only 60% of the available sunlight was able to penetrate the accumulated dust on the solar arrays. The MER Power subsystem operations team used the Multi-Mission Power Analysis Tool (MMPAT) to perform these Dust Factor calculations. The MMPAT software tool modeled the behavior of the solar arrays and the batteries as they interacted with the spacecraft power loads over the mission timeline. MMPAT also had knowledge (through telemetry and user inputs) of telemetered Power subsystem voltages and currents, atmospheric opacity (Tau), rover surface location, rover attitude, the planetary tilt and distance of Mars from the sun based on day of year, the instantaneous elevation of the sun based on time of day, temperatures (internal and external), terrain masking, and shadowing (due to the camera mast and antennas). Once all of the known sources of array energy loss are accounted, the remaining difference between the expected array energy and the actual array energy determines the solar array Dust Factor. The assumptions made while determining the Dust Factor are 1) that the single measured atmospheric opacity value (Tau) is constant over the course of the entire sol, 2) there is no measurable solar cell degradation, 3) there are no shorted strings and therefore, 4) all unexpected solar array energy losses are due to accumulated dust on the solar array. Although it cannot provide an absolute measure of dust loading, the determination of solar array Dust Factor provides a useful way of tracking dust accumulation and aeolian dust removal trends on Martian spacecraft. Any spacecraft on the Martian surface is vulnerable to dust, especially solar-powered spacecraft. The Spirit and Opportunity rovers were operational on the Martian surface for much longer than expected due in part to aeolian removal of dust from their solar arrays. The first few dust removal events were a pleasant surprise to the MER operations teams; however, over time a pattern began to arise. In over three Mars Years on the surface, the Power operations team tracked the solar array Dust Factor at Gusev Crater (location of MER-A) and observed that there were several significant dust removal and deposition events in Mars Year (MY) 27, even in the absence of a large dust storm. In MY 28 at Gusev Crater, the large atmospheric opacity (Tau) increase lagged significant dust removal events. Overall at Gusev Crater, there was a pattern of steady dust accumulation on the solar arrays, with a small number of significant dust cleaning events. At Meridiani Planum, where Opportunity rover operated for over seven Mars Years (late MY 26 to mid MY 34), a clear and consistent pattern of dust movement emerged. Meridiani Planum had a predictable, seasonally dependent pattern of gradual and continuous dust accumulation and removal. In summary, this paper explains the reasons for the development of the solar array Dust Factor and how it was used in mission operations. In particular, this paper describes the MMPAT software package, how it models array energy, including the important assumptions, model inputs and sources of error. And finally, this paper will show how the calculated solar array Dust Factor was used at Gusev Crater and Meridiani Planum to predict dust accumulation rates and weather patterns, and the importance of this generated data set for current and future solar- powered missions to Mars, such as the InSight lander and the planned Mars Sample Return rover.
A solar panel array with more than ten years space exposure was removed from the Mir core module in November 1997, and an eight panel section was returned to Earth in January 1998. Several solar cells were removed from panel eight of the returned array and placed in a high vacuum system with a residual gas analyzer (200 amu mass spectrometer) and a cold finger. Similar unflown solar cells of the same vintage were later obtained from Energia. Several of the unflown cells were also placed in the vacuum system and outgassed residues were collected on the LN2 cold finger. Almost 3 mg of outgassed residue was collected -from a string of three unflown solar cells over a period of 94 hours under vacuum. The collected residue was weighed with a microbalance, and then the residue was analyzed by FTIR spectroscopy, and by gas chromatograph-mass spectroscopy. About 25 outgassed constituents were separated by the gas chromatograph, and a high-resolution mass spectrum was obtained of each constituent. Molecular identifications have been made for the constituents. The constituents are primarily cyclic siloxanes, and several of the constituents are isomers of the same molecule. Most of the outgassed constituents have a molecular mass of about 500 amu. Almost one mg of residue was extracted from one sq cm of coverglass/adhesive from a flown solar cell by soaking in isopropyl alcohol for 30 minutes. The gas chromatograph separated about 20 constituents. The constituents are mostly cyclic siloxanes with linear branches, hydrocarbons, and phthalates. The typical molecular mass is about 600 amu. These identifications of specific outgassing molecules have resulted in a more complete understanding of the SiO(x) contamination on the Mir solar cell coverglasses, and on the MEEP experiment trays and optical specimens during the Shuttle-Mir Phase One flight experiment program. Adjusted outgassing rates based on the data reported here, and/or measured outgassing rates and specific molecular identifications of ISS hardware samples are needed to input into model predictions of induced environment effects of the ISS.
A battery charge regulator based on the series-connected boost regulator (SCBR) technology has been developed for high-voltage spacecraft applications. The SCBR regulates the solar array power during insolation to prevent battery overcharge or undercharge conditions. It can also be used to provide regulated battery output voltage to spacecraft loads if necessary. This technology uses industry-standard dc-dc converters and a unique interconnection to provide size, weight, efficiency, fault tolerance, and modularity benefits over existing systems. The high-voltage SCBR shown in the photograph has demonstrated power densities of over 1000 watts per kilogram (W/kg). Using four 150-W dc-dc converter modules, it can process 2500 W of power at 120 Vdc with a minimum input voltage of 90 Vdc. Efficiency of the SCBR was 94 to 98 percent over the entire operational range. Internally, the unit is made of two separate SCBR s, each with its own analog control circuitry, to demonstrate the modularity of the technology. The analog controllers regulate the output current and incorporate the output voltage limit with active current sharing between the two units. They also include voltage and current telemetry, on/off control, and baseplate temperature sensors. For peak power tracking, the SCBR was connected to a LabView-based data acquisition system for telemetry and control. A digital control algorithm for tracking the peak power point of a solar array was developed using the principle of matching the source impedance with the load impedance for maximum energy transfer. The algorithm was successfully demonstrated in a simulated spacecraft electrical system at the Boeing PhantomWorks High Voltage Test Facility in Seattle, Washington. The system consists of a 42-string, high-voltage solar array simulator, a 77-cell, 80-ampere-hour (A-hr) nickel-hydrogen battery, and a constant power-load module. The SCBR and the LabView control algorithm successfully tracked the solar array peak power point through various load transients, including sunlight discharge transients when the total load exceeded the maximum solar array output power.