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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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NASA NTRS2014

Impact of Acoustic Standing Waves on Structural Responses

For several decades large reverberant chambers and most recently direct field acoustic testing have been used in the aerospace industry to test larger structures with low surface densities such as solar arrays and reflectors to qualify them and to detect faults in the design and fabrication. It has been reported that in reverberant chamber and direct acoustic testing, standing acoustic modes may strongly couple with the fundamental structural modes of the test hardware (Reference 1). In this paper results from a recent reverberant chamber acoustic test of a composite reflector are discussed. These results provide further convincing evidence of the acoustic standing wave and structural modes coupling phenomenon. The purpose of this paper is to alert test organizations to this phenomenon so that they can account for the potential increase in structural responses and ensure that flight hardware undergoes safe testing. An understanding of the coupling phenomenon may also help minimize the over and/or under testing that could pose un-anticipated structural and flight qualification issues.

Reverberant acoustic field↗
NASA NTRS1985

Langley Research Center Photogrammetric Measurements of Solar Array Dynamics: Preliminary Results

The NASA Langley Research Center participated in the Solar Array Experiment with two primary objectives: (1) to study the structural and control dynamics of a new class of large, lightweight, low-frequency space structures, and (2) to develop technology for remote video measurement of structural motions. The shuttle orbiter's closed circuit television (CCTV) system was used to provide recorded video images of the solar array from four locations in the payload bay, two on the forward bulkhead and two on the aft bulkhead. White reflective targets were placed on the array to provide discrete points at which to track array motion. A dynamic test consisted of a quiescent period in which orbiter operations were inhibited and crew motion restricted, an excitation period, and a free decay period. The orbiter was placed in free drift while in a gravity gradient orientation and dynamics tests were timed to occur at orbital noon so that the Sun would illuminate one side of the array and Earth albedo the other. The CCTV system was turned on during the quiescent period, approximately 3 minutes prior to excitation. Video from all four CCTV cameras was recorded continuously through the excitation and free-decay periods for a total test record of 8 to 10 minutes.

Brumfield, M. L.↗
NASA NTRS1992

Influence of mass moment of inertia on normal modes of preloaded solar array mast

Earth-orbiting spacecraft often contain solar arrays or antennas supported by a preloaded mast. Because of weight and cost considerations, the structures supporting the spacecraft appendages are extremely light and flexible; therefore, it is vital to investigate the influence of all physical and structural parameters that may influence the dynamic behavior of the overall structure. The study primarily focuses on the mast for the space station solar arrays, but the formulations and the techniques developed in this study apply to any large and flexible mast in zero gravity. Furthermore, to determine the influence on the circular frequencies, the mass moment of inertia of the mast was incorporated into the governing equation of motion for bending. A finite element technique (MSC/NASTRAN) was used to verify the formulation. Results indicate that when the mast is relatively flexible and long, the mass moment inertia influences the circular frequencies.

Armand, Sasan C.↗
NASA NTRS2003

Thin Film Technology Development for the Powersphere

The Aerospace Corporation, NASA Glenn Research Center, Lockheed-Martin, and ILC Dover over the past two years have been engaged in developing a Multifunctional Inflatable Structure for the Powersphere Concept under contract with NASA (NAS3-01115). The Powersphere concept consists of a relatively large spherical solar array, which would be deployed from a microsatellite. The Powersphere structure and the deployment method was patented by the Aerospace Corporation (U.S. Patent Numbers 6,284,966 B 1 and 6,3 18,675). The work on this project has resulted in a number of technological innovations in the state of the art for integrating flexible thin-film solar cells with flex circuit harness technology and inflatable ultraviolet-light-rigidizable structures. The specific power, specific volume, for the Powersphere are presented in Figures 1 and 2 as a function of solar cell technology and efficiency. The Powersphere will enable microsatellite missions across NASA enterprises and DoD missions by providing ample electric power at an affordable cost. The Powersphere design provides attitude-independent electric power and thermal control for an enclosed microsatellite payload. The design is scalable, robust in high radiation environments and provides sufficient electric power to allow the use of electric propulsion. Electric propulsion enables precise positioning of microsatellites which is required for inspectors that would be deployed to inspect the International Space Station, Space Shuttle or large unmanned spacecraft. The Powersphere allows for efficient launch packaging versus deployed volume as shown in Figure 3.

Simburger, Edward J.↗
NASA NTRS1975

The attitude control of a solar electric propulsion /SEP/ spacecraft for an out-of-the-ecliptic mission

A computer-controlled attitude and articulation control system is proposed for a spacecraft of extremely soft structure that uses two 8-kW solar arrays with an area of nearly 100 sq m each to power six electric propulsion thrusters. The purpose of the articulation control is to point the scientific instruments, to orient the antenna toward the earth, and to reorient the thrust vector as required for trajectory control. The attitude control system provides control both during the low-thrust phase of the mission and during long-term spinning modes.

Marsh, E. L.↗
NASA NTRS1979

Extraterrestrial processing and manufacturing of large space systems. Volume 3: Executive summary

Facilities and equipment are defined for refining processes to commercial grade of lunar material that is delivered to a 'space manufacturing facility' in beneficiated, primary processed quality. The manufacturing facilities and the equipment for producing elements of large space systems from these materials and providing programmatic assessments of the concepts are also defined. In-space production processes of solar cells (by vapor deposition) and arrays, structures and joints, conduits, waveguides, RF equipment radiators, wire cables, converters, and others are described.

Miller, R. H.↗
NASA NTRS1986

A free-flying power plant for a manned space station

The 0.1 Hz fundamental structural frequency anticipated for the large solar array panels required to power manned space stations is close to the attitude control system bandwidth, and within the bandwidth of crew motion disturbances. Attention is presently given to a detachment of the solar panel system to create a free-flying powerplant for the space station, and to the control system that will maintain its attitude and position relative to the space station by means of thrusters that are supplied with space station state data by a laser/optical sensor system. Power transmission alternatives for this free-flying power system are discussed.

Paluszek, M.↗
NASA NTRS1989

Application of optical distributed sensing and computation to control of a large flexible spacecraft

Advances in real-time holography make possible the development of new distributed optical sensors and processors that may have application to the control of flexible structures. This paper presents the analytic evaluation of control system concepts utilizing this new technology on realistic spacecraft. The design of a distributed estimator based on this new optical sensing technique is targeted to vibration control of a complex spacecraft with multiple optical sensors, each of which views a portion of the spacecraft. A simulation of a complex truss structure that was flown in the Solar Array Flight Experiment is utilized to evaluate the performance of the estimator. It is shown that good performance can be obtained from a distributed estimator even in the presence of significant sensor noise.

Montgomery, R. C.↗
NASA NTRS1992

Electrical breakdown of Space Station Freedom surfaces

Space Station Freedom (SSF) will be the largest and highest power spacecraft that the U.S. has put into orbit. The solar array will generate 160 volts nominal when in sunlight, and the present baseline design is for the negative end of the solar array to be tied to SSF structure. Due to the balance of leakage currents through the plasma, the structure will be driven approximately 140 volts negative of the ambient conductive ionospheric plasma. Surface materials such as anodized aluminum will have this voltage drop across a thin dielectric which may not have sufficient dielectric strength to prevent dielectric breakdown. This can lead to arcing on the exterior surfaces of Space Station.

Carruth, M. R., Jr.↗
NASA NTRS2004

Thin-Film Solar Cells on Metal Foil Substrates for Space Power

Photovoltaic arrays have played a key role in power generation in space. The current technology will continue to evolve but is limited in the important mass specific power metric (MSP or power/weight ratio) because it is based on bulk crystal technology. The objective of this research is to continue development of an innovative photovoltaic technology for satellite power sources that could provide up to an order of magnitude saving in both weight and cost, and is inherently radiation-tolerant through use of thin film technology and thin foil substrates such as 5-mil thick stainless steel foil or 1-mil thick Ti. Current single crystal technology for space power can cost more than $300 per watt at the array level and weigh more than 1 kg/sq m equivalent to specific power of approx. 65 W/kg. Thin film material such as CuIn(1-x),Ga(x)S2, (CIGS2), CuIn(1-x), G(x)Se(2-y),S(y), (CIGSS) or amorphous hydrogenated silicon (a-Si:H) may be able to reduce both the cost and mass per unit area by an order of magnitude. Manufacturing costs for solar arrays are an important consideration for total spacecraft budget. For a medium sized 5kW satellite, for example, the array manufacturing cost alone may exceed $2 million. Moving to thin film technology could reduce this expense to less than $500 K. Previous work at FSEC demonstrated the potential of achieving higher efficiencies from CIGSS thin film solar cells on 5-mil thick stainless steel foil as well as initial stages of facility augmentation for depositing thin film solar cells on larger (6"x 4") substrates. This paper presents further progress in processing on metal foil substrates. Also, previous work at DayStar demonstrated the feasibility of flexible-thin-film copper-indium-gallium-diselenide (CIGS) solar cells with a power-to-weight ratio in excess of 1000 W/kg. We will comment on progress on the critical issue of scale-up of the solar cell absorber deposition process. Several important technical issues need to be resolved to realize the benefits of lightweight technologies for solar arrays, such as: monolithic interconnects, lightweight array structures, and new ultra-light support and deployment mechanisms. Once the technology has gained spaceflight certification it should find rapid acceptance in specific satellite markets.

Raffaelle, Ryne P.↗
NASA NTRS2004

Thin-Film Solar Cells on Polymer Substrates for Space Power

Photovoltaic arrays have played a key role in power generation in space. The current technology will continue to evolve but is limited in the important mass specific power metric (MSP or power/weight ratio) because it is based on bulk crystal technology. Solar cells based on thin-film materials offer the promise of much higher MSP and much lower cost. However, for many space applications, a 20% or greater AM0 efficiency (eta) may be required. The leading thin-film materials, amorphous Si, CuInSe, and CdTe have seen significant advances in efficiency over the last decade but will not achieve the required efficiency in the near future. Several new technologies are herein described to maximize both device eta and MSP. We will discuss these technologies in the context of space exploration and commercialization. One novel approach involves the use of very lightweight polyimide substrates. We describe efforts to enable this advance including materials processing and device fabrication and characterization. Another approach involves stacking two cells on top of each other. These tandem devices more effectively utilize solar radiation by passing through non-absorbed longer wavelength light to a narrow-bandgap bottom cell material. Modeling of current devices in tandem format indicates that AM0 efficiencies near 20% can be achieved with potential for 25% in the near future. Several important technical issues need to be resolved to realize the benefits of lightweight technologies for solar arrays, such as: monolithic interconnects, lightweight array structures, and new ultra-light support and deployment mechanisms. Recent advances will be stressed.

Hepps, A. F.↗
NASA NTRSDate not supplied

Manufacturing and Metal Extraction Lunar Technology (MMELT)

Extraction processes of oxygen from regolith under development include carbothermal, molten regolith electrolysis (MRE), ionic liquids, and others, all of which result in a metal slag byproduct. The separation of some of these metals from the slag byproduct have been accomplished on small scales. However, no system currently exists that can extract aluminum, from these processes at scale, to create suitable feedstocks for manufacturing. Funding for MMELT will be used to develop a demonstration of the first end-to-end process to additively manufacture aluminum wire from lunar regolith in a simulated environment.

Manufacturing↗
NASA NTRS2013

30-kW SEP Spacecraft as Secondary Payloads for Low-Cost Deep Space Science Missions

The Solar Array System contracts awarded by NASA's Space Technology Mission Directorate are developing solar arrays in the 30 kW to 50 kW power range (beginning of life at 1 AU) that have significantly higher specific powers (W/kg) and much smaller stowed volumes than conventional rigid-panel arrays. The successful development of these solar array technologies has the potential to enable new types of solar electric propulsion (SEP) vehicles and missions. This paper describes a 30-kW electric propulsion vehicle built into an EELV Secondary Payload Adapter (ESPA) ring. The system uses an ESPA ring as the primary structure and packages two 15-kW Megaflex solar array wings, two 14-kW Hall thrusters, a hydrazine Reaction Control Subsystem (RCS), 220 kg of xenon, 26 kg of hydrazine, and an avionics module that contains all of the rest of the spacecraft bus functions and the instrument suite. Direct-drive is used to maximize the propulsion subsystem efficiency and minimize the resulting waste heat and required radiator area. This is critical for packaging a high-power spacecraft into a very small volume. The fully-margined system dry mass would be approximately 1120 kg. This is not a small dry mass for a Discovery-class spacecraft, for example, the Dawn spacecraft dry mass was only about 750 kg. But the Dawn electric propulsion subsystem could process a maximum input power of 2.5 kW, and this spacecraft would process 28 kW, an increase of more than a factor of ten. With direct-drive the specific impulse would be limited to about 2,000 s assuming a nominal solar array output voltage of 300 V. The resulting spacecraft would have a beginning of life acceleration that is more than an order of magnitude greater than the Dawn spacecraft. Since the spacecraft would be built into an ESPA ring it could be launched as a secondary payload to a geosynchronous transfer orbit significantly reducing the launch costs for a planetary spacecraft. The SEP system would perform the escape from Earth and then the heliocentric transfer to the science target.

Dawn spacecraft↗
NASA NTRS1986

SAFE/DAE: Modal test in space

In September of 1984, NASA performed a series of experiments on orbit with a large solar wing attached to the Space Shuttle orbiter. These experiments, the Solar Array Flight Experiment (SAFE), mark the first tests of a large space structure in space. Extension, retraction, and dynamic tests had to be performed in space due to the fragility of the solar array. Due to the extendable and retractable design of the solar array, accelerometers and associated wires could not be used; therefore, remote sensing, the Dynamics Augmentation Experiment (DAE), was added to the SAFE program. The DAE uses a remote sensor based on star tracker technology to measure the dynamic response of the solar array. The DAE sensor tracked 18 targets on the solar array during free-decay response to a transient excitation. An overview of the SAFE/DAE is presented, highlighting analysis results from the remotely sensed data. Modal parameter estimates from the remotely sensed data were computed using the complex exponential and polyreference techniques.

Nesman, T. E.↗
NASA NTRS2018

SpiderFab: Process for On-Orbit Construction of Kilometer-Scale Apertures

The SpiderFab effort has investigated the value proposition and feasibility of radically changing the way we build and deploy spacecraft by enabling space systems to fabricate and integrate key components on-orbit. In this Phase II effort, we have focused on developing and demonstrating tools and processes to enable robotic systems to manufacture and assemble high performance structural elements that will serve as the support structures for components such as antennas and solar arrays. Through testing of these technologies in the laboratory environment,these efforts have established the technical feasibility of the key capabilities required for in-space manufacture of large apertures such as antennas, solar arrays, and optical systems,maturing prototype technical solutions for these capabilities to TRL-4. The SpiderFab effort has resulted in successful post-NIAC transition of the technology, first to SBIR-funded development of a technology for in-space manufacture (ISM) of truss structures, and then to a NASA/STMD Tipping Point Technologies funded effort to prepare a flight demonstration of ISM of a structure for a GEO communications satellite.

Manufacturing↗
NASA NTRS1995

Static stability of a three-dimensional space truss

In order to deploy large flexible space structures it is necessary to develop support systems that are strong and lightweight. The most recent example of this aerospace design need is vividly evident in the space station solar array assembly. In order to accommodate both weight limitations and strength performance criteria, ABLE Engineering has developed the Folding Articulating Square Truss (FASTMast) support structure. The FASTMast is a space truss/mechanism hybrid that can provide system support while adhering to stringent packaging demands. However, due to its slender nature and anticipated loading, stability characterization is a critical part of the design process. Furthermore, the dire consequences surely to result from a catastrophic instability quickly provide the motivation for careful examination of this problem. The fundamental components of the space station solar array system are the (1) solar array blanket system, (2) FASTMast support structure, and (3) mast canister assembly. The FASTMast once fully deployed from the canister will provide support to the solar array blankets. A unique feature of this structure is that the system responds linearly within a certain range of operating loads and nonlinearly when that range is exceeded. The source of nonlinear behavior in this case is due to a changing stiffness state resulting from an inability of diagonal members to resist applied loads. The principal objective of this study was to establish the failure modes involving instability of the FASTMast structure. Also of great interest during this effort was to establish a reliable analytical approach capable of effectively predicting critical values at which the mast becomes unstable. Due to the dual nature of structural response inherent to this problem, both linear and nonlinear analyses are required to characterize the mast in terms of stability. The approach employed herein is one that can be considered systematic in nature. The analysis begins with one and two-dimensional failure models of the system and its important components. From knowledge gained through preliminary analyses a foundation is developed for three-dimensional analyses of the FASTMast structure. The three-dimensional finite element (FE) analysis presented here involves a FASTMast system one-tenth the size of the actual flight unit. Although this study does not yield failure analysis results that apply directly to the flight article, it does establish a method by which the full-scale mast can be evaluated.

Shaker, John F.↗
NASA NTRS1995

Cost Trade Between Multi-Junction, Gallium Arsenide, and Silicon Solar Cells

Multi-junction (MJ), gallium arsenide (GaAs), and silicon (Si) solar cells have respective test efficiencies of approximately 24%, 18.5% and 14.8%. Multi-junction and gallium arsenide solar cells weigh more than silicon solar 2 cells and cost approximately five times as much per unit power at the cell level. A trade is performed for the TRMM spacecraft to determine which of these cell types would have offered an overall performance and price advantage to the spacecraft. A trade is also performed for the multi-junction cells under the assumption that they will cost over ten times that of silicon cells at the cell level. The trade shows that the TRMM project, less the cost of the instrument, ground systems and mission operations, would spend approximately $552,000 dollars per kilogram to launch and suppon3science in the case of the spacecraft equipped with silicon solar cells. If these cells are changed out for gallium arsenide solar cells, an additional 31 kilograms of science can be launched and serviced at a price of approximately $90 thousand per kilogram. The weight reduction is shown to derive from the smaller area of the array and hence reductions in the weight of the array substrate and supporting structure. ff the silicon solar cells are changed out for multi-junction solar cells, an additional 45 kilograms of science above the silicon base line can be launched and supported at a price of approximately $58,000 per kilogram. The trade shows that even if the multi-junction cells are priced over ten times that of silicon cells, a price that is much higher than projected, that the additional 45 kilograms of science are launched and serviced at $180,000 per kilogram. This is still much less than the original $552,000 per kilogram to launch and service the science. Data and qualitative factors are presented to show that these figures are subject to a great deal of uncertainty. Nonetheless, the benefit of the higher efficiency solar cells for TRMM is far greater than the uncertainties in the analysis.

Gaddy, Edward M.↗
NASA NTRS1988

Oxidation and protection of fiberglass-epoxy composite masts for photovoltaic arrays in the low Earth orbital environment

Fiberglass-epoxy composites are considered for use as structural members for the mast of the space station solar array panel. The low Earth orbital environment in which space station is to operate is composed mainly of atomic oxygen, which has been shown to cause erosion of many organic materials and some metals. Ground based testing in a plasma asher was performed to determine the extent of degradation of fiberglass-epoxy composites when exposed to a simulated atomic oxygen environment. During exposure, the epoxy at the surface of the composite was oxidized, exposing individual glass fibers which could easily be removed. Several methods of protecting the composite were evaluated in an atomic oxygen environment and with thermal cycling and flexing. The protection techniques evaluated to date include an aluminum braid covering, an indium-tin eutectic and a silicone based paint. The open aluminum braid offered little protection while the CV-1144 coating offered some initial protection against atomic oxygen, but appears to develop cracks which accelerate degradation when flexed. Coatings such as the In-Sn eutectic may provide adequate protection by containing the glass fibers even though mass loss still occurs.

Rutledge, Sharon K.↗