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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 NTRS2018

Lessons Learned in the Flight Qualification of the S-NPP and NOAA-20 Solar Array Mechanisms

Deployable solar arrays are the energy source used on almost all Earth orbiting spacecraft and their release and deployment are mission-critical; fully testing them on the ground is a challenging endeavor. The 8 meter long deployable arrays flown on two sequential NASA weather satellites were each comprised of three rigid panels almost 2 meters wide. These large panels were deployed by hinges comprised of stacked constant force springs, eddy current dampers, and were restrained through launch by a set of four releasable hold-downs using shape memory alloy release devices. The ground qualification testing of such unwieldy deployable solar arrays, whose design was optimized for orbital operations, proved to be quite challenging and provides numerous lessons learned. A paperwork review and follow-up inspection after hardware storage determined that there were negative torque margins and missing lubricant, this paper will explain how these unexpected issues were overcome. The paper will also provide details on how the hinge subassemblies, the fully-assembled array, and mechanical ground support equipment were subsequently improved and qualified for a follow-on flight with considerably less difficulty. The solar arrays built by Ball Aerospace Corp. for the Suomi National Polar Partnership (S-NPP) satellite and the Joint Polar Satellite System (JPSS-1) satellite (now NOAA-20) were both successfully deployed on-obit and are performing well.

Helfrich, Daniel↗
NASA NTRS2018

Lessons Learned in the Flight Qualification of the S-NPP and NOAA-20 Solar Array Mechanisms

Deployable solar arrays are the energy source used on almost all Earth orbiting spacecraft and their release and deployment are mission-critical; fully testing them on the ground is a challenging endeavor. The 8 meter long deployable arrays flown on two sequential NASA weather satellites were each comprised of three rigid panels almost 2 meters wide. These large panels were deployed by hinges comprised of stacked constant force springs, eddy current dampers, and were restrained through launch by a set of four releasable hold-downs using shape memory alloy release devices. The ground qualification testing of such unwieldy deployable solar arrays, whose design was optimized for orbital operations, proved to be quite challenging and provides numerous lessons learned. A paperwork review and follow-up inspection after hardware storage determined that there were negative torque margins and missing lubricant, this paper will explain how these unexpected issues were overcome. The paper will also provide details on how the hinge subassemblies, the fully-assembled array, and mechanical ground support equipment were subsequently improved and qualified for a follow-on flight with considerably less difficulty. The solar arrays built by Ball Aerospace Corp. for the Suomi National Polar Partnership (SNPP) satellite and the Joint Polar Satellite System (JPSS-1) satellite (now NOAA-20) were both successfully deployed on-obit and are performing well.

Sexton, Adam↗
NASA NTRS1972

The 110 watt per kilogram lightweight solar array system

This ultralightweight solar array system is applied to three generic mission types: (1) interplanetary, (2) geosynchronous, and (3) manned space station. The requirements of each of these missions, as they pertain to the solar array, are presented. Existing lightweight solar array system concepts are reviewed, along with conclusions regarding the applicability of this technology to the feasibility of the ultralightweight solar array system. Several system concepts are included for further evaluation. The existing technology base, as it pertains to solar cells, solar cell covers, interconnects and substrates, and deployable booms, is reviewed. The attitude control of spacecraft with large flexible solar arrays is also discussed.

Shepard, N. F.↗
NASA NTRS1993

Solar array electrical performance assessment for Space Station Freedom

Electrical power for Space Station Freedom will be generated by large Photovoltaic arrays with a beginning of life power requirement of 30.8 kW per array. The solar arrays will operate in a Low Earth Orbit (LEO) over a design life of fifteen years. This paper provides an analysis of the predicted solar array electrical performance over the design life and presents a summary of supporting analysis and test data for the assigned model parameters and performance loss factors. Each model parameter and loss factor is assessed based upon program requirements, component analysis, and test data to date. A description of the LMSC performance model, future test plans, and predicted performance ranges are also given.

Smith, Bryan K.↗
NASA NTRS1993

Solar array electrical performance assessment for Space Station Freedom

Electrical power for Space Station Freedom will be generated by large photovoltaic arrays with a beginning of life power requirement of 30.8 kW per array. The solar arrays will operate in a Low Earth Orbit (LEO) over a design life of fifteen years. This paper provides an analysis of the predicted solar array electrical performance over the design life and presents a summary of supporting analysis and test data for the assigned model parameters and performance loss factors. Each model parameter and loss factor is assessed based upon program requirements, component analysis and test data to date. A description of the LMSC performance model future test plans and predicted performance ranges are also given.

Smith, Bryan K.↗
NASA NTRS1996

The ARA Mark 3 solar array design and development

The ARA (Advanced Rigid Array) Mark3 solar array of Fokker Space BV is currently in its final stages of qualification (wing tests to be completed in March, 1996; unit/part tests in April, 1996). With regard to its predecessor, the ARA Mark2, the design has not only been improved in terms of mechanical and electrical performance, but also with regard to production cost and throughput time. This 'state of the art' array is designed to fit the needs of a wide variety of geostationary telecommunications satellites and is qualified for launch on the complete range of medium/large size commercial launchers (Ariane IV & V, Atlas, Delta, Proton, Long March, H2). The first mission to fly the new ARA Mk3 array is Hot Bird 2 (customer: Eutelsat, prime contractor: Matra Marconi Space; launch: mid-1996). In this configuration, its end of life (EOL) power-to-mass ratio is 42 W/kg, with an operational life of more than 12 years. The main mechanisms on a solar array are typically found in the deployment system and in the hold down and release system. During the design and development phase of these mechanisms, extensive engineering and qualification tests have been performed. This paper presents the key design features of these mechanisms and the improvements that were made with regard to their predecessors. It also describes the qualification philosophy on unit/part and wing level. Finally, some of the development items that turned out to be critical, as well as the lessons learned from them, are discussed.

vanHassel, Rob H. A.↗
NASA NTRS1999

Retrieval of Mir Solar Array

A Russian solar array panel removed in November 1997 from the non-articulating photovoltaic array on the Mir core module was returned to Earth on STS-89 in January 1998. The panel had been exposed to low Earth orbit (LEO) for 10 years prior to retrieval. The retrieval provided a unique opportunity to study the effects of the LEO environment on a functional solar array. To take advantage of this opportunity, a team composed of members from RSC-Energia (Russia), the Boeing Company, and the following NASA Centers--Johnson Space Center, Kennedy Space Center, Langley Research Center, Marshall Space Flight Center, and Lewis Research Center--was put together to analyze the array. After post-retrieval inspections at the Spacehab Facility at Kennedy in Florida, the array was shipped to Lewis in Cleveland for electrical performance tests, closeup photodocumentation, and removal of selected solar cells and blanket material. With approval from RSC-Energia, five cell pairs and their accompanying blanket and mesh material, and samples of painted handrail materials were selected for removal on the basis of their ability to provide degradation information. Sites were selected that provided different sizes and shapes of micrometeoroid impacts and different levels of surface contamination. These materials were then distributed among the team for round robin testing.

Rutledge, Sharon K.↗
NASA NTRS1994

Plasma chamber testing of advanced photovoltaic solar array coupons

The solar array module plasma interactions experiment is a space shuttle experiment designed to investigate and quantify the high voltage plasma interactions. One of the objectives of the experiment is to test the performance of the Advanced Photovoltaic Solar Array (APSA). The material properties of array blanket are also studied as electric insulators for APSA arrays in high voltage conditions. Three twelve cell prototype coupons of silicon cells were constructed and tested in a space simulation chamber.

TRT-EXPERIMENTAL↗
NASA NTRS1989

Effect of micrometeorid and space debris impacts on the Space Station Freedom solar array surfaces

Both solar and antisolar surfaces of the Space Station Freedom solar arrays are vulnerable to micrometeoroid and space debris impacts. Impacts on the solar surface result in damage to the active area of the solar cell and a corresponding reduction in generated power. Impacts on the antisolar surface could result in damage to the circuit which interconnects the cells which in turn may produce open circuit strings or panels. An evaluation is presented of the power degradation resulting from the impacts of micrometeoroid and space debris on the solar surface of the array blanket. Moreover, given a particle diameter that could damage the circuit interconnecting the cells, the probability of an open circuit panel is computed, along with the probability that the solar array blanket will meet the power requirement over the design lifetime.

Nahra, Henry K.↗
NASA NTRS2008

Electromagnetically Clean Solar Arrays

The term 'electromagnetically clean solar array' ('EMCSA') refers to a panel that contains a planar array of solar photovoltaic cells and that, in comparison with a functionally equivalent solar-array panel of a type heretofore used on spacecraft, (1) exhibits less electromagnetic interferences to and from other nearby electrical and electronic equipment and (2) can be manufactured at lower cost. The reduction of electromagnetic interferences is effected through a combination of (1) electrically conductive, electrically grounded shielding and (2) reduction of areas of current loops (in order to reduce magnetic moments). The reduction of cost is effected by designing the array to be fabricated as a more nearly unitary structure, using fewer components and fewer process steps. Although EMCSAs were conceived primarily for use on spacecraft they are also potentially advantageous for terrestrial applications in which there are requirements to limit electromagnetic interference. In a conventional solar panel of the type meant to be supplanted by an EMCSA panel, the wiring is normally located on the back side, separated from the cells, thereby giving rise to current loops having significant areas and, consequently, significant magnetic moments. Current-loop geometries are chosen in an effort to balance opposing magnetic moments to limit far-0field magnetic interactions, but the relatively large distances separating current loops makes full cancellation of magnetic fields problematic. The panel is assembled from bare photovoltaic cells by means of multiple sensitive process steps that contribute significantly to cost, especially if electomagnetic cleanliness is desired. The steps include applying a cover glass and electrical-interconnect-cell (CIC) sub-assemble, connecting the CIC subassemblies into strings of series-connected cells, laying down and adhesively bonding the strings onto a panel structure that has been made in a separate multi-step process, and mounting the wiring on the back of the panel. Each step increases the potential for occurrence of latent defects, loss of process control, and attrition of components. An EMCSA panel includes an integral cover made from a transparent material. The silicone cover supplants the individual cover glasses on the cells and serves as an additional unitary structural support that offers the advantage, relative to glass, of the robust, forgiving nature of the silcone material. The cover contains pockets that hold the solar cells in place during the lamination process. The cover is coated with indium tin oxide to make its surface electrically conductive, so that it serves as a contiguous, electrically grounded shield over the entire panel surface. The cells are mounted in proximity to metallic printed wiring. The painted-wiring layer comprises metal-film traces on a sheet of Kapton (or equivalent) polyimide. The traces include contact pads on one side of the sheet for interconnecting the cells. Return leads are on the opposite side of the sheet, positioned to form the return currents substantially as mirror images of, and in proximity to, the cell sheet currents, thereby minimizing magnetic moments. The printed-wiring arrangement mimics the back-wiring arrangement of conventional solar arrays, but the current-loop areas and the resulting magnetic moments are much smaller because the return-current paths are much closer to the solar-cell sheet currents. The contact pads are prepared with solder fo electrical and mechanical bonding to the cells. The pocketed cover/shield, the solar cells, the printed-wiring layer, an electrical bonding agent, a mechanical-bonding agent, a composite structural front-side face sheet, an aluminum honeycomb core, and a composite back-side face sheet are all assembled, then contact pads are soldered to the cells and the agents are cured in a single lamination process.

Stem, Theodore G.↗
NASA NTRS2020

On Orbit Flight Testing of the Roll-Out Solar Array

The Roll-Out Solar Array (ROSA) is an innovative, lightweight solar array with a flexible substrate that makes use of the stored strain energy in its composite structural members to provide deployment without the use of motors while still packaging efficiently. The ROSA flight experiment was launched to the International Space Station (ISS)on June 3rd, 2017 as a part of the eleventh SpaceX commercial resupply mission. After two weeks of storage, ROSA was extracted from the Dragon spacecraft, deployed, and put through a week of tests to verify its structural characteristics as well as the performance of photovoltaics. The structural dynamics experiments included active excitation of out-of-plane bending modes, and a study of the thermal-structural interaction of the structure during day-night transitions. Data was gathered from a small number of accelerometers located at key points on the solar array and from cameras observing numerous photogrammetry targets spread out over the structure and the photovoltaic blanket. The goal in this effort was to better understand the performance of ROSA and to improve modeling efforts for future designs of similar solar arrays. Of particular interest are the first few system modes and mode shapes of the array, the amount of structural damping present, and degree of structural-thermal interaction seen during eclipse exit.

ROSA↗
NASA NTRS2015

Thin, Flexible IMM Solar Array

NASA needs solar arrays that are thin, flexible, and highly efficient; package compactly for launch; and deploy into large, structurally stable high-power generators. Inverted metamorphic multijunction (IMM) solar cells can enable these arrays, but integration of this thin crystalline cell technology presents certain challenges. The Thin Hybrid Interconnected Solar Array (THINS) technology allows robust and reliable integration of IMM cells into a flexible blanket comprising standardized modules engineered for easy production. The modules support the IMM cell by using multifunctional materials for structural stability, shielding, coefficient of thermal expansion (CTE) stress relief, and integrated thermal and electrical functions. The design approach includes total encapsulation, which benefits high voltage as well as electrostatic performance.

Walmsley, Nicholas↗
NASA NTRS2002

SMEX-Lite Modular Solar Array Architecture

For the most part, Goddard solar arrays have been custom designs that are unique to each mission. The solar panel design has been frozen prior to issuing an RFP for their procurement. There has typically been 6-9 months between RFP release and contract award, followed by an additional 24 months for performance of the contract. For Small Explorer (SMEX) missions, with three years between mission definition and launch, this has been a significant problem. The SMEX solar panels have been sufficiently small that the contract performance period has been reduced to 12-15 months. The bulk of this time is used up in the final design definition and fabrication of flight solar cell assemblies. Even so, it has been virtually impossible to have the spacecraft design at a level of maturity sufficient to freeze the solar panel geometry and release the RFP in time to avoid schedule problems with integrating the solar panels to the spacecraft. With that in mind, the SMEX-Lite project team developed a modular architecture for the assembly of solar arrays to greatly reduce the cost and schedule associated with the development of a mission- specific solar array. In the modular architecture, solar cells are fabricated onto small substrate panels. This modular panel (approximately 8.5" x 17" in this case) becomes the building block for constructing solar arrays for multiple missions with varying power requirements and geometrical arrangements. The mechanical framework that holds these modules together as a solar array is the only mission-unique design, changing in size and shape as required for each mission. There are several advantages to this approach. First, the typical solar array development cycle requires a mission unique design, procurement, and qualification including a custom qualification panel. With the modular architecture, a single qualification of the SMEX-Lite modules and the associated mechanical framework in a typical configuration provided a qualification by similarity to multiple missions. It then becomes possible to procure solar array modules in advance of mission definition and respond quickly and inexpensively to a selected mission's unique requirements. The solar array modular architecture allows the procurement of solar array modules before the array geometry has been frozen. This reduces the effect of procurement lead-time on the mission integration and test flow by as much as 50%. Second, by spreading the non-recurring costs over multiple missions, the cost per unit area is also reduced. In the case of the SMEX-Lite procurement, this reduction was by about one third of the cost per unit area compared to previous SMEX mission-unique procurements. Third, the modular architecture greatly facilitates the infusion of new solar cell technologies into flight programs as these technologies become available. New solar cell technologies need only be fabricated onto a standard-sized module to be incorporated into the next available mission. The modular solar array can be flown in a mixed configuration with some new and some standard cell technologies. Since each module has its own wiring terminals, the array can be arranged as desired electrically with little impact to cost and schedule. The solar array modular architecture does impose some additional constraints on systems and subsystem engineers. First, they must work with discrete solar array modules rather than size the array to fit exactly within an available envelope. The array area is constrained to an integer multiple of the module area. Second, the modular design is optimized for space radiation and thermal environments not greatly different from a typical SMEX LEO environment. For example, a mission with a highly elliptical orbit (e.g., Polar, SMEX/FAST) would require thicker coverglasses to protect the solar cells from the more intense radiation environment.

Lyons, John↗
NASA NTRS2016

Micrometeorite Impact Test of Flex Solar Array Coupon

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.

Wright, K. H.↗
NASA NTRS1984

STS 41-D Solar Array Flight Experiment

The Solar Array Flight Experiment (SAFE) developed under the direction of the Marshall Space Flight Center, scheduled for launch on STS 41-D, will demonstrate a lightweight solar array technology which offers a factor of 3 improvement in weight and a factor of 10 improvement in specific volume over solar array systems currently in use in the space program. The experiment, which will include multiple deployment and retraction demonstration, verification of electrical and thermal performance, and verification of structural dynamic math models is 15 feet by 105 feet in size and, if completely covered with solar cells, would produce approximately 12.5 kW of electrical power. The unit has now been developed, tested, and is at Kennedy Space Center (KSC) being prepared for launch.

Turner, G. F.↗
NASA NTRS2016

Micrometeorite Impact Test of Flex Solar Array Coupon

Spacecraft with solar arrays operate throughout the near earth environment and are planned for outer planet missions. An often overlooked test condition for solar arrays that is applicable to these missions is micrometeoroid impacts and possibly electrostatic discharge (ESD) events resulting from these impacts. NASA Marshall Space Flight Center (MSFC) is partnering with Space Systems/Loral, LLC (SSL) to examine the results of simulated micrometeoroid impacts on the electrical performance of an advanced, lightweight flexible solar array design. The test is performed at MSFC's Micro Light Gas Gun Facility with SSL-provided coupons. Multiple impacts were induced at various locations on a powered test coupon under different string voltage (0V-150V) and string current (1.1A - 1.65A) conditions. The setup, checkout, and results from the impact testing are discussed.

Wright, K. H.↗