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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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Top level overviews of the PILS and VSAT programs to be presented to the NASA Engineering Safety Center (NESC) Electrical Power Technology Discipline Team Annual Meeting
Technology and analytical techniques for development of solar power system for space stations
Solar array welding technology is examined from its beginnings in the late 1960's to the present. The U.S. and European efforts are compared, and significant similarities are highlighted. The utilization of welding technology for space use is shown to have been influenced by a number of subtle, secondary factors.
The results of all the major program phases of the program are reported. All goals of the program, which are listed were successfully accomplished and are briefly described. A complete list is included of all drawings generated during this program.
The current space PV (photovoltaic) technology development program of ESA is described. The program is closely coupled to the European space mission scenario for the next 10 year period and has as its main objective to make the most effective use of the limited resources available for technology in the present economical climate. This requires a well-balanced approach between concentration on very few options and keeping the competition alive if more than one promising technology exists. The paper describes ESA's main activities in the areas of solar array technology, solar cell technology, solar cell assembly technology, and special test and verification activities including the in-orbit demonstration of new technologies.
NASA is currently developing advanced solar cell and solar array technologies to support future exploration activities. These advanced photovoltaic technology development efforts are needed to enable very large (multi-hundred kilowatt) power systems that must be compatible with solar electric propulsion (SEP) missions. The technology being developed must address a wide variety of requirements and cover the necessary advances in solar cell, blanket integration, and large solar array structures that are needed for this class of missions. Th is paper will summarize NASA's plans for high power SEP missions, initi al mission studies and power system requirements, plans for advanced photovoltaic technology development, and the status of specific cell and array technology development and testing that have already been conducted.
The solar array system considered is composed of two wings. Each wing consists of a solar array blanket, a blanket launch storage container, an extension/retraction mast assembly, a blanket tensioning system, and an array electrical harness. A technology evaluation is performed to assess the applicable solar array state-of-the-art and to define the supporting effort necessary to achieve technology readiness for meeting the Solar Electric Propulsion Stage (SEPS) solar array design requirements. Details of mechanical design are discussed along with questions of electrical design, operational reliability advantages, and array assembly advantages.
Many innovations in spacecraft power and propulsion have been recently tested at NASA, particularly in non-chemical propulsion. One improvement in solar array technology is solar concentration using thin polymer film Fresnel lenses. Weight and cost savings were proven with the Solar Concentrator Arrays with Refractive Linear Element Technology (SCARLET)-II array on NASA's Deep Space I spacecraft. The Fresnel lens concentrates solar energy onto high-efficiency solar cells, decreasing the area of solar cells needed for power. Continued efficiency of this power system relies on the thin film's durability in the space environment and maintaining transmission in the 300 - 1000 nm bandwidth. Various polymeric materials have been tested for use in solar concentrators, including Lexan(TM), polyethylene terephthalate (PET), several formulations of Tefzel(Tm) and Teflon(TM), and DC 93-500, the material selected for SCARLET-II. Also tested were several innovative materials including Langley Research Center's CPI and CP2 polymers and atomic oxygen- resistant polymers developed by Triton Systems, Inc. The Environmental Effects Group of the Marshall Space Flight Center's Materials, Processes, and Manufacturing Department exposed these materials to simulated space environment and evaluated them for any change in optical transmission. Samples were exposed to a minimum of 1000 equivalent Sun hours of near-UV radiation (250 - 400 nm wavelength). Materials that appeared robust after near-UV exposure were then exposed to charged particle radiation equivalent to a five-year dose in geosynchronous orbit. These exposures were performed in MSFC's Combined Environmental Effects Test Chamber, a unique facility with the capability to expose materials simultaneously or sequentially to protons, low-energy electrons, high-energy electrons, near UV radiation and vacuum UV radiation. Reflectance measurements can be made on the samples in vacuum. Prolonged exposure to the space environment will decrease the polymer film's transmission and thus reduce the conversion efficiency. A method was developed to normalize the transmission loss and thus rank the materials according to their tolerance to space environmental exposure. Spectral results and the material ranking according to transmission loss are presented.
The solar arrays flight experiment consists of four experiments on two different flights. The first experiment, termed the baseline, has a basic purpose to demonstrate the flight readiness of lightweight solar array technology for solar electric propulsion and other payload power applications. The early availability of this experiment and its basic large space structure characteristics make it a logical candidate to demonstrate other disciplines critical to large space structures. These demonstrations form the basis for three other solar array experiments, two in remote sensing and one in control. All of these experiments are briefly reviewed in this paper.
The technology developments required and a preliminary design of a lightweight 25 kW solar array for the solar electric propulsion stage (SEPS) have been defined. The requirements for a 65 W/Kg SEPS solar array system requires significant component weight reductions over present state-of-the-art flexible solar arrays in both electrical and structural-mechanical designs. A requirement for operation from 0.3 au to 6.0 au presents a wide range of temperature environments as well as severe combined thermal/vacuum/UV radiation environments. Additional requirements are capability for partial array retraction operation, and capability for full retraction and automatic preloading for survival of the Shuttle reentry environment. An assessment of current lightweight flexible solar array technology is made against the SEPS solar array requirements and new technology requirements are defined. A preliminary design and the operating characteristics of a flat-fold solar array system meeting the SEPS requirements is presented. A full-width, 10-ft-tall functional array model, including representative welded electrical modules and a model astromast, was fabricated and tested.
In 1985, the Jet Propulsion Laboratory initiated the Advanced Photovoltaic Solar Array (APSA) program. The program objective is to demonstrate a producible array system by the early 1990s with a specific performance of at least 130 W/kG (beginning-of-life) as an intermediate milestone towards the long range goal of 300 W/kG. The APSA performance represents an approximately four-fold improvement over existing rigid array technology and a doubling of the performance of the first generation NASA/OAST SAFE flexible blanket array of the early 1980s.
The SCARLET I (Solar Concentrator Army with Refractive Linear Element Technology) solar array wing was designed and built to demonstrate, in flight, the feasibility of integrating deployable concentrator optics within the design envelope of typical rigid array technology. Innovative mechanism designs were used throughout the array, and a full series of qualification tests were successfully performed in anticipation of a flight on the Multiple Experiment Transporter to Earth Orbit and Return (METEOR) spacecraft. Even though the Conestoga launch vehicle was unable to place the spacecraft in orbit, the program effort was successful in achieving the milestones of analytical and design development functional validation, and flight qualification, thus leading to a future flight evaluation for the SCARLET technology.
Two special instruments were custom built for our critical study and manufactured and employed to test the SA materials of interest to cryogenic temperature conditions. We have procured custom manufactured two independent systems such as Thermo -Mechanical Analyzer (TMA) with Three Point Bend Probe (TPBP) module and Dilatometer. TMA can be used to measure Young’ s modulus or elastic modulus (e -modulus). Dilatometer can be used to measure Co -efficient of Thermal Expansion (CTE). We have optimized the materials configuration to measure physical properties of radiated and non -radiated materials for CTE and e -modulus, which will be useful to assess materials and processes reliability for cryogenic temperature applications. We have successfully implemented for some materials that are of interest to solar panels for the planned ECM project to measure CTE and Young's modulus. Also a separate study was undertaken to thermal cycle qualify SAs for ECM project. One can do thermal cycling to temperatures of 50 K and 133 K using liquid Helium (Lq He). This is in reality exorbitantly too expensive and is also inefficient. Furthermore, this is not technically controllable test to 133 K and 50 K temperatures. This is also too expensive to accomplish. Therefore, a vacuum set -up was made along with a cryostat successfully. This is a closed loop system where we do not lose the Helium gas. We were able to get to 35 K for a given test coupon size. These temperatures are lower than what we would intend to qualify the solar array technologies. This will allow to have some margin at cold temperatures. We successfully completed the qualification of solar array technologies down to 50 K temperatures. This is the first time we are reporting this accomplishment. In this paper, we will present the solar array materials properties for cryogenic temperatures down to 20 K temperatures and the thermal cycling qualification test results from 133 K to 50 K for three times the mission life of 120 thermal cycles to meet JPL Design.
Engineering design of a near-term laser surface-to-surface power beaming station to transmit power to users in shadowed regions or during the night. Input system requirements were to be able to provide 300 W of continuous usable power to users including landers and rovers, at distances up to 10 km, with a total system landed mass under 625 kg. Due to surface irregularities and the close horizon of the moon, to achieve 10-km transmission the laser must be elevated above the surface. The Vertical Solar Array Technology (VSAT) program is developing a solar array mounted on a 10-m tall mast, intended to fit on a Commercial Lunar Payload Services (CLPS) lander, with target readiness date of 2028. At an optimum location near the south pole, the elevated array produces power for a majority of the lunar day.
A baseline configuration for a three section folding array, with retraction capability, was developed which would utilize electroformed aluminum hollowcore substrates and beryllium frames. The three section array was not fabricated because of difficulties with impurities in the aluminum electroforming bath. A procedure was developed for etching the copper mandrel from virtually any size of aluminum hollowcore panel in approximately one hour. Procedures were developed for analyzing the content of peroxide, water, total aluminum, and lithium-aluminum-hydride in an aluminum electroforming solution.
The development of ultrathin silicon solar cell array modules from initial design to flight testing is discussed. Three 80-cell modules were subjected to the thermal soak test, the LEO thermal cycle test, and the solar array flight experiment, and six 48-cell welded modules were evaluated in the geosynchronous orbit thermal cycle test. It is observed that the electrical performance of the modules was not affected by the different environmental conditions. The automatic assembly of the cell modules, in particular the welding and solar cell glassing operation, is described. The specific power capabilities of Kapton, Kapton-Kevlar-Kapton, Kapton-graphite-Kapton, and Kapton-graphite-aluminum honeycomb-graphite solar array designs are assessed.
Silicon, aluminum, and glass are the primary raw materials that will be required for production of solar arrays on the moon. A process sequence is proposed for producing these materials from lunar regolith is proposed, consisting of separating the required materials from lunar rock with fluorine. Fluorosilane produced by this process is reduced to silicon; the fluorine salts are reduced to metals by reaction with metallic potassium. Fluorine is recovered from residual MgF and CaF2 by reaction with K2O. Aluminum, calcium oxide, and magnesium oxide are recovered to manufacture structural materials and glass.
Using a laser to send power to a photovoltaic receiver has been proposed to transmit electrical power on the moon, particularly for applications such as powering a rover in near-polar permanent-ly-shadowed regions (PSR) where solar power is not available, however detailed engineering design studies of the spacecraft for such applications has not previously been undertaken. In this work, we did an engineering design study of two applications of laser power beaming for near-term lunar ap-plications. In the first application studied, an application was studied to provide power during the lunar night for a global network of small landers spread across the lunar surface, at latitudes ranging from equatorial to near polar landing sites. We analyze a proposal to power such small landers from orbit, using a laser to beam power from an orbital power station to photovoltaic arrays on the landers that are tuned to the laser wavelength. A commercially-available high-power 1.07-µ diode-pumped fiber laser was chosen as the source. To minimize beam spread of the a spot at the required distance, a 1.5 meter optical element was provided, using a design based on the Kepler telescope. To provide power to landers at any location, three orbital stations (“beamcraft) are required, each carrying a 3-kW laser. (If the surface science platforms requiring power are only in polar locations, or converse-ly, only in near equatorial locations, only one orbital platform is needed.) The approach is seen to be feasible, and a systems analysis was completed, the concept of operations for the system out-lined and a design for the beamcraft put together. The second design study looked at surface-to-surface power beaming using the VSAT as the la-ser platform, for an application to provide power to the interior of a permanently shadowed lunar crater from a surface platform. To maximize the distance of beaming, taking into account possible surface irregularities and the short distance to the horizon of the moon, it is desirable to emplace the laser at an elevation above the surface. The Vertical Solar Array Technology (VSAT) is a NASA program developing a solar array mounted vertically on a 10-m tall mast, designed for emplacement on a Commercial Lunar Payload Services (CLPS) lander to provide 10-kW (BOL) power near the south polar region of the moon, with a target readiness date of 2028. We used this design as the starting platform and the power source for a laser power beaming station. By mounting the laser beam director at the top of the solar array mast, a viewing distance to power receivers up to 10 km is possible. Requirements for the system were to be able to provide 300 W of continuous usable power to users including CLPS landers, VIPER class rovers, or the proposed Lunar Terrain Vehicle. The specified requirement was to be able to transmit power to a distance of up to 10 km, over a design lifetime of 5 years, and fitting within a total system landed mass under 625 kg. Again, a 1.07-µ fiber laser is mounted on the deck of the lander, with laser output sent to the laser beam director by a fiber-optic cable. A 7 square meter deployable radiator keeps the laser within operating temperature limits. The beam director is based on the design of a prototype unit developed by the University of California Santa Barbara. The system beams power for 57% of the time, with 44% of the time idle (accounting for the time when the VSAT array is itself in shadow). 1595 Watts of optical power are output in the beam. Accounting for receiver efficiency and beam losses, this results in an output onto the 1.5-meter receiving photovoltaic array of 542 watts. Of this, 300 watts is directly available to the user, while 242 watts is directed to the batteries for use while the beam is not available.