Search NASA⌕ Search

SEARCH · Search NASA

Results for “Solar Array Structures”

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.

At least 127 records · Page 7
All sourcesNASA NTRSDOE OSTI
NASA NTRS2012

Feasibility of Large High-Powered Solar Electric Propulsion Vehicles: Issues and Solutions

Human exploration beyond low Earth orbit will require the use of enabling technologies that are efficient, affordable, and reliable. Solar electric propulsion (SEP) has been proposed by NASA s Human Exploration Framework Team as an option to achieve human exploration missions to near Earth objects (NEOs) because of its favorable mass efficiency as compared to traditional chemical systems. This paper describes the unique challenges and technology hurdles associated with developing a large high-power SEP vehicle. A subsystem level breakdown of factors contributing to the feasibility of SEP as a platform for future exploration missions to NEOs is presented including overall mission feasibility, trip time variables, propellant management issues, solar array power generation, array structure issues, and other areas that warrant investment in additional technology or engineering development.

Capadona, Lynn A.↗
NASA NTRS2011

Atomic Oxygen Fluence Monitor

This innovation enables a means for actively measuring atomic oxygen fluence (accumulated atoms of atomic oxygen per area) that has impinged upon spacecraft surfaces. Telemetered data from the device provides spacecraft designers, researchers, and mission managers with real-time measurement of atomic oxygen fluence, which is useful for prediction of the durability of spacecraft materials and components. The innovation is a compact fluence measuring device that allows in-space measurement and transmittance of measured atomic oxygen fluence as a function of time based on atomic oxygen erosion yields (the erosion yield of a material is the volume of material that is oxidized per incident oxygen atom) of materials that have been measured in low Earth orbit. It has a linear electrical response to atomic oxygen fluence, and is capable of measuring high atomic oxygen fluences (up to >10(exp 22) atoms/sq cm), which are representative of multi-year low-Earth orbital missions (such as the International Space Station). The durability or remaining structural lifetime of solar arrays that consist of polymer blankets on which the solar cells are attached can be predicted if one knows the atomic oxygen fluence that the solar array blanket has been exposed to. In addition, numerous organizations that launch space experiments into low-Earth orbit want to know the accumulated atomic oxygen fluence that their materials or components have been exposed to. The device is based on the erosion yield of pyrolytic graphite. It uses two 12deg inclined wedges of graphite that are over a grit-blasted fused silica window covering a photodiode. As the wedges erode, a greater area of solar illumination reaches the photodiode. A reference photodiode is also used that receives unobstructed solar illumination and is oriented in the same direction as the pyrolytic graphite covered photodiode. The short-circuit current from the photodiodes is measured and either sent to an onboard data logger, or transmitted to a receiving station on Earth. By comparison of the short-circuit currents from the fluence-measuring photodiode and the reference photodiode, one can compute the accumulated atomic oxygen fluence arriving in the direction that the fluence monitor is pointing. The device produces a signal that is linear with atomic oxygen fluence using a material whose atomic oxygen erosion yield has been measured over a period of several years in low-Earth orbit.

Banks, Bruce A.↗
NASA NTRS2015

An Analytical Solution for Yaw Maneuver Optimization on the International Space Station and Other Orbiting Space Vehicles

This paper suggests a new method for optimizing yaw maneuvers on the International Space Station (ISS). Yaw rotations are the most common large maneuvers on the ISS often used for docking and undocking operations, as well as for other activities. When maneuver optimization is used, large maneuvers, which were performed on thrusters, could be performed either using control moment gyroscopes (CMG), or with significantly reduced thruster firings. Maneuver optimization helps to save expensive propellant and reduce structural loads ‐ an important factor for the ISS service life. In addition, optimized maneuvers reduce contamination of the critical elements of the vehicle structure, such as solar arrays. This paper presents an analytical solution for optimizing yaw attitude maneuvers. Equations describing pitch and roll motion needed to counteract the major torques during a yaw maneuver are obtained. A yaw rate profile is proposed. Also the paper describes the physical basis of the suggested optimization approach. In the obtained optimized case, the torques are significantly reduced. This torque reduction was compared to the existing optimization method which utilizes the computational solution. It was shown that the attitude profiles and the torque reduction have a good match for these two methods of optimization. The simulations using the ISS flight software showed similar propellant consumption for both methods. The analytical solution proposed in this paper has major benefits with respect to computational approach. In contrast to the current computational solution, which only can be calculated on the ground, the analytical solution does not require extensive computational resources, and can be implemented in the onboard software, thus, making the maneuver execution automatic. The automatic maneuver significantly simplifies the operations and, if necessary, allows to perform a maneuver without communication with the ground. It also reduces the probability of command errors. The suggested analytical solution provides a new method of maneuver optimization which is less complicated, automatic and more universal. A maneuver optimization approach, presented in this paper, can be used not only for the ISS, but for other orbiting space vehicles.

Dobrinskaya, Tatiana↗
NASA NTRS1977

Solar photovoltaic power stations

The subsystems of a solar photovoltaic central power system are identified and the cost of major components are estimated. The central power system, which would have a peak power capability in the range of 50 to 1000 MW, utilizes two types of subsystems - a power conditioner and a solar array. Despite differences in costs of inverters, the overall cost of the total power conditioning subsystem is about the same for all approaches considered. A combination of two inverters operating from balanced dc buses as a pair of 6-pulse groups is recommended. A number of different solar cell modules and tracking array structures were analyzed. It is concluded that when solar cell costs are high (greater than $500/kW), high concentration modules are more cost effective than those with low concentration. Vertical-axis tracking is the most effective of the studied tracking modes. For less expensive solar cells (less than $400/kW), fixed tilt collector/reflector modules are more cost effective than those which track.

Chowaniec, C. R.↗
NASA NTRS2004

A 100 kW-Class Technology Demonstrator for Space Solar Power

A first step in the development of solar power from space is the flight demonstration of critical technologies. These fundamental technologies include efficient solar power collection and generation, power management and distribution, and thermal management. In addition, the integration and utilization of these technologies into a viable satellite bus could provide an energy-rich platform for a portfolio of payload experiments such as wireless power transmission (WPT). This paper presents the preliminary design of a concept for a 100 kW-class free-flying platform suitable for flight demonstration of technology experiments. Recent space solar power (SSP) studies by NASA have taken a stepping stones approach that lead to the gigawatt systems necessary to cost-effectively deliver power from space. These steps start with a 100 kW-class satellite, leading to a 500 kW and then a 1 MW-class platform. Later steps develop a 100 MW bus that could eventually lead to a 1-2 GW pilot plant for SSP. Our studies have shown that a modular approach is cost effective. Modular designs include individual laser-power-beaming satellites that fly in constellations or that are autonomously assembled into larger structures at geosynchronous orbit (GEO). Microwave power-beamed approaches are also modularized into large numbers of identical units of solar arrays, power converters, or supporting structures for arrays and microwave transmitting antennas. A cost-effective approach to launching these modular units is to use existing Earth-to-orbit (ETO) launch systems, in which the modules are dropped into low Earth orbit (LEO) and then the modules perform their own orbit transfer to GEO using expendable solar arrays to power solar electric thrusters. At GEO, the modules either rendezvous and are assembled robotically into larger platforms, or are deployed into constellations of identical laser power-beaming satellites. Since solar electric propulsion by the modules is cost-effective for both self-transport of the modules from LEO to GEO, and for on-orbit stationkeeping and repositioning capability during the satellite’s lifetime, this technology is also critical in technology development for SSP. The 100 kW-class technology demonstrator will utilize advanced solar power collection and generation technologies, power management and distribution, advanced thermal management, and solar electric propulsion. State-of-the-art solar concentrators, highly efficient multi-junction solar cells, integrated thermal management on the arrays, and innovative deployable structure design and packaging make the 100 kW satellite feasible for launch on one existing launch vehicle. Early SSP studies showed that a major percentage of the on-orbit mass for power-beaming satellites was from massive power converters at the solar arrays, at the bus, at the power transmitter, or at combinations of these locations. Higher voltage mays and power management and distribution (PMAD) systems reduce or eliminate the need for many of these massive power converters, and could enable direct-drive of high-voltage solar electric thrusters. Lightweight, highly efficient thermal management systems are a critical technology that must be developed and flown for SSP feasibility. Large amounts of power on satellites imply that large amounts of waste heat will need to be managed. In addition, several of the more innovative lightweight configurations proposed for SSP satellites take advantage of solar concentrators that are intractable without advanced thermal management technologies for the solar arrays. These thermal management systems include efficient interfaces with the WPT systems or other high-power technology experiments, lightweight deployable radiators that can be easily integrated into satellite buses, and efficient reliable thermal distribution systems that can pipe heat from the technology experiments to the radiators. In addition to demonstrating the integration and use of these mission-critical technologies, the 100 kw-class satellite will provide a large experiment deck for a portfolio of technology experiments. Current plans for this technology demonstrator allow 2000 kg of payload capability and up to 100 kW of power. The technology experiments could include one or more wireless power transmission demonstrations, either to the Earth’s surface or to a suitable space-based receiver. Technology experiments to quantify the on-orbit performance of critical technologies for SSP or space exploration are welcomed. In addition, the technology experiments provide an opportunity for international cooperation, to advance technology readiness levels of SSP technologies that require flight demonstration. This paper will present the preliminary design for a 100 kW solar-powered satellite and a variety of technology experiments that may be suitable for flight demonstration. In addition, a space-to-Earth-surface WPT experiment will be discussed.

Connie Carrington↗
NASA NTRS1991

Durability evaluation of photovoltaic blanket materials exposed on LDEF tray S1003

Several candidate protective coatings on Kapton and uncoated Kapton were exposed to the LEO environment on the LDEF in order to determine whether the coatings could be used to protect polymeric substrates from degradation in the LEO environment. These materials are used for flexible solar array panels in which the polymer is the structural member that supports the solar cell and current carriers. Arrays such as these are used on the Hubble Space Telescope and will be used on Space Station Freedom. The results of the experiments are presented.

Rutledge, Sharon K.↗
NASA NTRS2012

FAST Mast Structural Response to Axial Loading: Modeling and Verification

The International Space Station s solar array wing mast shadowing problem is the focus of this paper. A building-block approach to modeling and analysis is pursued for the primary structural components of the solar array wing mast structure. Starting with an ANSYS (Registered Trademark) finite element model, a verified MSC.Nastran (Trademark) model is established for a single longeron. This finite element model translation requires the conversion of several modeling and analysis features for the two structural analysis tools to produce comparable results for the single-longeron configuration. The model is then reconciled using test data. The resulting MSC.Nastran (Trademark) model is then extended to a single-bay configuration and verified using single-bay test data. Conversion of the MSC. Nastran (Trademark) single-bay model to Abaqus (Trademark) is also performed to simulate the elastic-plastic longeron buckling response of the single bay prior to folding.

Knight, Norman F., Jr.↗
NASA NTRS1992

Grounding of space structures

Space structures, such as the Space Station solar arrays, must be extremely light-weight, flexible structures. Accurate prediction of the natural frequencies and mode shapes is essential for determining the structural adequacy of components, and designing a controls system. The tension pre-load in the 'blanket' of photovoltaic solar collectors, and the free/free boundary conditions of a structure in space, causes serious reservations on the use of standard finite element techniques of solution. In particular, a phenomenon known as 'grounding', or false stiffening, of the stiffness matrix occurs during rigid body rotation. This paper examines the grounding phenomenon in detail. Numerous stiffness matrices developed by others are examined for rigid body rotation capability, and found lacking. A force imbalance inherent in the formulations examined is the likely cause of the grounding problem, suggesting the need for a directed force formulation.

Bosela, P. A.↗
NASA NTRS1992

A new pre-loaded beam geometric stiffness matrix with full rigid body capabilities

Space structures, such as the Space Station solar arrays, must be extremely light-weight, flexible structures. Accurate prediction of the natural frequencies and mode shapes is essential for determining the structural adequacy of components, and designing a controls system. The tension pre-load in the 'blanket' of photovoltaic solar collectors, and the free/free boundary conditions of a structure in space, causes serious reservations on the use of standard finite element techniques of solution. In particular, a phenomenon known as 'grounding', or false stiffening, of the stiffness matrix occurs during rigid body rotation. The authors have previously shown that the grounding phenomenon is caused by a lack of rigid body rotational capability, and is typical in beam geometric stiffness matrices formulated by others, including those which contain higher order effects. The cause of the problem was identified as the force imbalance inherent in the formulations. In this paper, the authors develop a beam geometric stiffness matrix for a directed force problem, and show that the resultant global stiffness matrix contains complete rigid body mode capabilities, and performs very well in the diagonalization methodology customarily used in dynamic analysis.

Bosela, P. A.↗
NASA NTRS2001

ISS And Space Environment Interactions Without Operating Plasma Contactor

The International Space Station (ISS) will be the largest, highest power spacecraft placed in orbit. Because of this the design of the electrical power system diverged markedly from previous systems. The solar arrays will operate at 160 V and the power distribution voltage will be 120 V. The structure is grounded to the negative side of the solar arrays so under the right circumstances it is possible to drive the ISS potential very negative. A plasma contactor has been added to the ISS to provide control of the ISS structure potential relative to the ambient plasma. The ISS requirement is that the ISS structure not be greater than 40 V positive or negative of local plasma. What are the ramifications of operating large structures with such high voltage power systems? The application of a plasma contactor on ISS controls the potential between the structure and the local plasma, preventing degrading effects. It is conceivable that there can be situations where the plasma contactor might be non-functional. This might be due to lack of power, the need to turn it off during some of the build-up sequences, the loss of functionality for both plasma contactors before a replacement can be installed, similar circumstances. A study was undertaken to understand how important it is to have the contactor functioning and how long it might be off before unacceptable degradation to ISS could occur. The details of interaction effects on spacecraft have not been addressed until driven by design. This was true for ISS. If the structure is allowed to float highly negative impinging ions can sputter exposed conductors which can degrade the primary surface and also generate contamination due to the sputtered material. Arcing has been known to occur on solar arrays that float negative of the ambient plasma. This can also generate electromagnetic interference and voltage transients. Much of the ISS structure and pressure module surfaces exposed to space is anodized aluminum. The anodization thickness is very thin to provide the required solar absorptance and emittance. For conditions where ISS structure can charge negative a large percentage of the array voltage, the dielectric strength of this layer is low, and dielectric breakdown (arcing) can occur. The energy stored capacitively in the structure can be delivered to the arc. The mechanisms by which this energy is delivered and how much of the energy is available hasn't been fully quantified. Questions have been raised regarding the possibility of whether a sustained arc might result due to current collected by the solar arrays from local plasma. It was postulated that even if dielectric breakdown didn't occur, impacts due to micrometeoroids and space debris could penetrate thin layers of dielectric on ISS and initiate an arc due to the coupling provided by the dense local plasma produced by the impact. This was proven in experiments conducted jointly by MSFC and Auburn University. A target chamber with a simulated ionospheric plasma and a biased, anodized aluminum plate and a 1-microfarad capacitor was used. The plate was then impacted by 75-micron particles accelerated to orbital velocity. Arc discharges were sustained for higher voltages but a threshold appears below which no discharge was initiated. Most items without an exposed power system will float electrically near the local plasma potential. This is true of the Space Shuttle, an Astronaut on EVA, and similar items. The structure of ISS might be at a large negative voltage. Therefore, capacitively stored energy can be transferred during docking, installing external boxes and equipment and Astronaut contact with ISS structure. The circumstances of when this can happen and the resulting effects are evaluated in this study. Also, a crewmember on EVA might be in the vicinity of an arc. All safety aspects of such an encounter including charging, molten particles from the arc site and EMI have been evaluated. This paper will report on the total results of this study focussed on the 4A configuration, scheduled to be complete in November, 2000. Interactions such as arcing, debris induced arcs, sustained arcs, sputtering, contamination from sputtering and arcing, docking interactions and Astronaut safety issues will all be addressed.

Carruth, M. R., Jr.↗
NASA NTRS1973

Solar Electric Propulsion Stage design

This paper summarizes the analyses and design studies performed for the Solar Electric Propulsion Stage (SEPS). Mission analysis results were used to establish general design requirements. A performance sensitivity analysis was then used to define an operational power level of 21 kW for accomplishing a broad range of interplanetary and earth-orbital missions. With this base, propulsion system and stage subsystem design analyses were performed to support concept and configuration studies. Design study areas include solar array, thrust subsystem, thermal control, structures, and astrionics. Results are presented primarily for areas not covered by more specific technical papers (i.e., thrust subsystem, solar array, design requirements and configuration studies).

Masek, T. D.↗
NASA NTRS1978

A high volume cost efficient production macrostructuring process

The paper presents an experimentally developed surface macro-structuring process suitable for high volume production of silicon solar cells. The process lends itself easily to automation for high throughput to meet low-cost solar array goals. The tetrahedron structure observed is 0.5 - 12 micron high. The surface has minimal pitting with virtually no or very few undeveloped areas across the surface. This process has been developed for (100) oriented as cut silicon. Chemi-etched, hydrophobic and lapped surfaces were successfully texturized. A cost analysis as per Samics is presented.

Chitre, S. R.↗
NASA NTRS1985

NASA-VCOSS dynamic test facility

The Large Space Structure Ground Test Facility under development at the NASA Marshall Space Flight Center in Huntsville, Alabama is described. The Ground Test Facility was established initially to test experimentally the control system to be used on the Solar Array flight Experiment. The structural dynamics of the selected test article were investigated, including the fidelity of the associated mathematical model. The facility must permit the investigation of structural dynamics phenomena and be able to evaluate candidate attitude control and vibration suppression techniques.

Waites, H. B.↗
NASA NTRS1991

The effect of atomic oxygen on polysiloxane-polyimide for spacecraft applications in low Earth orbit

Polysiloxane-polyimide films are of interest as a replacement for polyimide Kapton in the Space Station Freedom solar array blanket. The blanket provides the structural support for the solar cells as well as providing transport of heat away from the back of the cells. Polyimide Kapton would be an ideal material to use; however, its high rate of degradation due to attack by atomic oxygen in low Earth orbit, at the altitudes Space Station Freedom will fly, is of such magnitude that if left unprotected, the blanket will undergo structural failure in much less than the desired 15 year operating life. Polysiloxane-polyimide is of interest as a replacement material because it should from its own protective silicon dioxide coating upon exposure to atomic oxygen. Mass, optical, and photomicrographic data obtained in the evaluation of the durability of polysiloxane-polyimide to an atomic oxygen environment are presented.

Rutledge, Sharon K.↗
NASA NTRS1988

Assessment of the effects of space debris and meteoroids environment on the space station solar array assembly

The methodology used to assess the probability of no impact of space debris and meteoroids on a spacecraft structure is applied to the Space Station solar array assembly. Starting with the space debris and meteoroids flux models, the projected surface area of the solar cell string circuit of the solar array panel and the mast longeron, and the design lifetime, the possibility of no impact on the solar array mast and solar cell string circuits was determined as a function of particle size. The probability of no impact on the cell string circuits was used to derive the probability of no open circuit panel. The probability of meeting a certain power requirement at the end of the design lifetime was then calculated as a function of impacting particle size. Coupled with a penetration and damage models/correlations which relate the particle size to the penetration depth and damage, the results of this analysis can be used to determine the probability of meeting the lower power requirement given a degree of redundancy, and the probability of no impact on the solar array mast.

Nahra, Henry K.↗
NASA NTRS1988

Assessment of the effects of space debris and meteoroids environment on the Space Station solar array assembly

The methodology used to assess the probability of no impact of space debris and meteoroids on a spacecraft structure is applied to the Space Station solar array assembly. Starting with space debris and meteoroid flux models, the projected surface area of the solar cell string circuit of the solar array panel and the mast longeron, and the design lifetime, the possibility of no impact on the solar mast and solar cell string circuits was determined as a function of particle size. The probability of no impact on the cell string circuits was used to derive the probability of no open circuit panel. The probability of meeting a certain power requirement at the end of the design lifetime was then calculated as a function of impacting particle size. Coupled with a penetration and damage models/correlations which relate the particle size to penetration depth and damage, the results of this analysis can be used to determine the probability of meeting the lower power requirement given a degree of redundancy, and the probability of no impact on the solar array mast.

Nahra, Henry K.↗