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 613 records · Page 34
All sourcesNASA NTRSDOE OSTI
NASA NTRS1982

Low concentration ratio solar array for low Earth orbit multi-100 kW application

An ongoing preliminary design effort directed toward a low-concentration-ratio photovoltaic array system based on 1984 technology and capable of delivering multi-hundred kilowatts (300 kW to 1000 kW range) in low earth orbit is described. The array system consists of two or more array modules each capable of delivering between 80 kW to 172 kW using silicon solar cells or gallium arsenide solar cells respectively. The array module deployed area is 1320 square meters and consists of 4356 pryamidal concentrator elements. The module, when stowed in the Space Shuttle's payload bay, has a stowage volume of a cube with 3.24 meters on a side. The concentrator elements are sized for a geometric concentration ratio (GCR) of six with an aperture area of 0.5 meters x 0.5 meters. The structural analysis and design trades leading to the baseline design are discussed. The configuration, as well as optical, thermal and electrical performance analyses that support the design and overall performance estimates for the array are described.

Nalbandian, S. J.↗
NASA NTRS1988

High resolution VLA-Nancay observations of the sun

The Very Large Array (VLA) and the Nancay Radioheliograph (NR) can be used in the 1990s to provide unique observations of the quiescent and active corona. The combined data can specify the evolution and magnetic structure at a variety of heights in the quiescent corona, from compact to extended structures such as loops, streamers and coronal holes. The scientific potential of simultaneous VLA-NR research is illustrated by observations of solar noise storms with hitherto unavailable resolution in space (VLA) and time (NR). Unique information will also be provided for other types of solar bursts, leading to an improved understanding of solar activity during the coming maximum.

Kerdraon, A.↗
NASA NTRS2014

High-Resolution Submillimeter and Near-Infrared Studies of the Transition Disk Around Sz 91

To reveal the structures of a transition disk around a young stellar object in Lupus, Sz 91, we have performed aperture synthesis 345 GHz continuum and CO(3--2) observations with the Submillimeter Array (approximately 1" - 3" resolution), and high-resolution imaging of polarized intensity at the K(sub s) -band by using the HiCIAO instrument on the Subaru Telescope (0.25" resolution). Our observations successfully resolved the inner and outer radii of the dust disk to be 65 AU and 170 AU, respectively, which indicates that Sz 91 is a transition disk source with one of the largest known inner holes. The model fitting analysis of the spectral energy distribution reveals an H 2 mass of 2.4×10(exp −3) M(solar mass) in the cold (T less than 30 K) outer part at 65 less than r less than 170 AU by assuming a canonical gas-to-dust mass ratio of 100, although a small amount (greater than 3×10(exp −9) M(solar mass)) of hot (T approximately 180 K) dust possibly remains inside the inner hole of the disk. The structure of the hot component could be interpreted as either an unresolved self-luminous companion body (not directly detected in our observations) or a narrow ring inside the inner hole. Significant CO(3--2) emission with a velocity gradient along the major axis of the dust disk is concentrated on the Sz 91 position, suggesting a rotating gas disk with a radius of 420 AU. The Sz 91 disk is possibly a rare disk in an evolutionary stage immediately after the formation of protoplanets because of the large inner hole and the lower disk mass than other transition disks studied thus far.

canonical gas-to-dust mass ratio↗
NASA NTRS1977

Design of satellite flexibility experiments

A preliminary study has been completed to begin development of a flight experiment to measure spacecraft control/flexible structure interaction. The work reported consists of two phases: identification of appropriate structural parameters which can be associated with flexibility phenomena, and suggestions for the development of an experiment for a satellite configuration typical of near-future vehicles which are sensitive to such effects. Recommendations are made with respect to the type of data to be collected and instrumentation associated with these data. The approach consists of developing the equations of motion for a vehicle possessing a flexible solar array, then linearizing about some nominal motion of the craft. A set of solutions are assumed for array deflection using a continuous normal mode method and important parameters are exposed. Inflight and ground based measurements are distinguished. Interrelationships between these parameters, measurement techniques, and input requirements are discussed which assure minimization of special vehicle maneuvers and optimization of data to be obtained during the normal flight sequence.

Kaplan, M. H.↗
NASA NTRS2018

Pterodactyl: Integrated Control Design for Precision Targeting of Deployable Entry Vehicles

Deployable Entry Vehicles (DEVs) enable in-situ scientific exploration at destinations with atmospheres across the solar system. Because they stow in a compact form and deploy only when ready to enter the atmosphere, DEVs relax the volume constraint imposed by rigid aeroshells. This work seeks to do for a DEV what the Wright Brothers did to propel modern day aviation: develop the guidance and control (G&C) methods that will make maneuvering and precision landing of DEV a reality. The Pterodactyl project objective is to deliver an integrated G&C methodology for a DEV, based on a detailed analysis that utilizes a Multi-disciplinary, Design, Analysis and Optimization (MDAO) framework. The current state-of-the-art for blunt body entry, G&C is rooted in the precision landing of vehicles such as Mars Science Laboratory (MSL) and Apollo, which used a propulsive reaction control system (RCS) to steer. Recent research has taken a particular interest in non-propulsive control for DEVs, including direct force control (angle of attack modulation via control surfaces or mass movement) and drag modulation (discrete change in ballistic coefficient). Using the MDAO framework that includes a guidance and control model to explore multiple control concepts for a DEV will shed light on the best design approach for these vehicles. In Pterodactyl, we will complete this study for a novel DEV concept, and then we will fabricate a functional prototype to help validate the design. The project is expected to down-select to a final control architecture by the end of 2018, and complete fabrication of the prototype by the end of 2019.The DEV chosen for detailed study in this project is the Adaptable Deployable Entry and Placement Technology (ADEPT). ADEPT uses a revolutionary 3D-woven carbon fabric that is foldable, can serve as primary structure, and can survive the extreme heating environment of atmospheric entry. The specific configuration of ADEPT under investigation is called Lifting Nano-ADEPT (LNA). LNA is designed for secondary payloads missions that require precision landing either for scientific objectives at a target destination or for payload recovery at Earth.The MDAO framework being created through this research, called COBRA-Pt (Composite Beam Roll-Up Solar Array-Prototype), will combine three critical elements of the system design: a guidance algorithm with Monte Carlo, a parametric control model, and vehicle geometry details. Novel control models being studied are deployable aerodynamic surfaces as well as shape morphing. These concepts will be compared at the system level with a more traditional propulsive RCS by comparing several key performance parameters. Upon completion of the design study, a functional prototype of LNA will be fabricated that will include the integration of guidance software and relevant control actuators. We expect this study will provide critical data that could feed into the development of an Earth-based flight test of LNA. The COBRA-Pt framework will provide a modular system by which to study any DEV concept in any atmosphere.

Smith, Brandon↗
NASA NTRS1989

The experiments of LIPS 3

LIPS 3 is a member of the Living Plume Shield series of spacecraft. In each LIPS project, the plume shield, a simple sheet metal cone, was structurally stiffened, and an active satellite was then built around it. The original purpose of the plume shield was to prevent the plume from solid propellent engines, which are fired outside the atmosphere after the aerodynamic shroud is jettisoned, from reaching the primary payload. The surface of LIPS 3 facing the plume also functioned in this manner, but the anterior surfaces were unaffected, and it was there that all solar arrays, sensors, and experiments were mounted. The purpose of LIPS 3 was to provide a test bed for new space power sources. With the long delays projected for schedules of the STS and other major launch systems, it appeared that a decade might pass before long term flight data could be obtained on many new and innovative power sources. The fact that a launch scheduled for early in 1987 required a plume shield was seen as a unique opportunity to obtain some of this data in a timely manner. The LIPS 3 system, the experiments placed aboard, and the experiment data acquisition subsystem are described. Various problems were encountered during integration and after launch; those which appear to effect the accuracy of experimental results are discussed. A preliminary description is given of the accuracy of the flight experiment data.

Severns, J. G.↗
NASA NTRS2009

Lunar Reconnaissance Orbiter (LRO) Rapid Thermal Design Development

The Lunar Reconnaissance Orbiter (LRO) project had a rapid development schedule starting with project conception in spring of 2004, instrument and launch vehicle selection late in 2005 and then launch in early 2009. The lunar thermal environment is one of the harshest in our solar system with the heavy infrared loading of the moon due to low albedo, lack of lunar atmosphere, and low effective regolith conduction. This set of constraints required a thermal design which maximized performance (minimized radiator area and cold control heater power) and minimized thermal hardware build at the orbiter level (blanketing, and heater service). The orbiter design located most of the avionics on an isothermalized heat pipe panel called the IsoThermal Panel (ITP). The ITP was coupled by dual bore heat pipes to an Optical Solar Reflector (OSR) covered heat pipe radiator. By coupling all of the avionics to one system, the hardware was simplified. The seven instruments were mainly heritage instruments which resulted in their desired radiators being located by their heritage design. This minimized instrument redesigns and therefore allowed them to be delivered earlier, though it resulted in a more complex orbiter level blanket and heater service design. Three of the instruments were mounted on a tight pointing M55J optical bench that needed to be covered in heaters to maintain pointing. Two were mounted to spacecraft controlled radiators. One was mounted to the ITP Dual Bores. The last was mounted directly to the bus structure on the moon facing panel. The propulsion system utilized four-20 pound insertion thrusters and eight-5 pound attitude control thrusters (ACS) in addition to 1000 kg of fuel in two large tanks. The propulsion system had a heater cylinder and a heated mounting deck for the insertion thrusters which coupled most of the propulsion design together simplifying the heater design. The High Gain Antenna System (HGAS) and Solar Array System (SAS) used dual axis actuator gimbal systems. HGAS required additional boom heaters to cool the approximately 10 W of RF losses thru the rotary joints and wave guides from the 40 W Ka system. By design this module needed a fair amount of heater, blanketing, and radiator complexity. The SAS system required a separate cable wrap radiator to help cool the Solar Array harness which dissipated 30 W thru the actuators and cable wraps. This module also was complex.

Baker, Charles↗
NASA NTRS1973

Pioneer Venus mission plan for atmospheric probes and an orbiter.

The mission concept is described for a spin-stabilized spacecraft being designed to deliver an array of aerodynamic probes over the surface of Venus, and, on a second flight mission, to inject the spacecraft into a Venusian orbit. Large and Small Probes and the instrumented 'Bus' structure will enter the Venusian atmosphere and directly measure its characteristics. An orbiting configuration of the same basic Bus spacecraft will serve as a platform from which to examine the planetary environment, the outer atmosphere and its interaction with the solar wind, as well as surface topography. The orbital mission will extend over a complete revolution of the planet around the sun and around its own axis. Requirements for cost effectiveness and firm adherence within early cost projections are emphatic program guidelines which affect every major element of the mission plan and conceptual design.

Dyer, J. W.↗
NASA NTRS1976

Long Duration Exposure Facility - A multipurpose free-flying experiment carrier

The Long Duration Exposure Facility (LDEF) is being developed by NASA to accommodate, using the Space Shuttle, a class of technology, science and applications experiments which require a free-flying exposure in space, but which do not require a sophisticated or dedicated satellite. The facility is a simple reusable structure about 14 ft in diameter and 30 ft long, which is three-axis gravity-gradient stabilized when free flying in space. Various aspects of the LDEF are examined, including the attitude stabilization system, the experiment trays, the experiment exposure control canister,and the environment monitoring system. LDEF mission plans are considered, as is experiment selection - the space debris damage experiment, the thermal coating experiment, the solar cells and arrays experiment, and the active optical components experiment.

Kinard, W. H.↗
NASA NTRS1984

Wind loading on solar concentrators: Some general considerations

A survey was completed to examine the problems and complications arising from wind loading on solar concentrators. Wind loading is site specific and has an important bearing on the design, cost, performance, operation and maintenance, safety, survival, and replacement of solar collecting systems. Emphasis herein is on paraboloidal, two-axis tracking systems. Thermal receiver problems also are discussed. Wind characteristics are discussed from a general point of view. Current methods for determining design wind speed are reviewed. Aerodynamic coefficients are defined and illustrative examples are presented. Wind tunnel testing is discussed, and environmental wind tunnels are reviewed. Recent results on heliostat arrays are reviewed as well. Aeroelasticity in relation to structural design is discussed briefly.

Roschke, E. J.↗
NASA NTRS1985

Box truss development and its application

Since 1977, Martin Marietta Denver Aerospace has aggressively pursued development of deployable structural systems applicable to a wide variety of Shuttle-transportable large space system requirements. This effort has focused on the deployable box truss, mechanisms and materials development, mesh reflector design and fabrication, gate frame truss design and fabrication, and offset-fed antenna design and analysis. The activities discussed are: box truss design; metal matrix composites; precision joints; enhanced passive damping design; mesh reflector development; gate frame truss for solar arrays; 15-meter spinning radio meter; and 60 x 120 meter push broom antenna.

Coyner, J. V.↗
NASA NTRS1988

Relative planetary radar sensitivities: Arecibo and Goldstone

The increase of the Deep Space Network antennas from 64 meter to 70 meter diameter represents the first of several improvements that will be made over the next decade to enhance earth based radar sensitivity to solar system targets. The aperture increase at the Goldstone DSS-14 site, coupled with a proposed increase in transmitter power to 1000 kW, will improve the 3.5 cm radar by about one order of magnitude. Similarly, proposed Arecibo Observatory upgrades of a Gregorian feed structure and an increase of transmitter power to 1000 kW will increase the sensitivity of this radar about 20 fold. In addition, a Goldstone to Very Large Array bistatic observation with horizon to horizon tracking will have 3.5 times more sensitivity than will a Goldstone horizon to horizon monostatic observation. All of these improvements, which should be in place within the next decade, will enrich an already fertile field of planetary exploration.

Renzetti, N. A.↗
DOE OSTI2024

Solar Design/Install Labor Quantification (Cooperative Research and Development Final Report)

The goals of this project were to quantify the rate and/or total expected labor expenditure of engineers designing solar arrays and electricians installing solar across a spectrum of project sizes and configurations. Specifically, procuring and analyzing labor data from solar engineering firms relating to the amount of time it takes an engineer to design a solar project, the amount of time it takes an engineer to perform a wiring or stringing design, the amount of time it takes an electrician to install a solar project and the amount of time it takes electricians to install the strings of a solar array. Then these details could be analyzed to evaluate labor intensity for solar design/installation. These goals were unable to be completed because the participant was unable to collect the necessary data for analysis.

14 SOLAR ENERGY↗
NASA NTRS1986

High voltage system: Plasma interaction summary

The possible interactions that could exist between a high voltage system and the space plasma environment are reviewed. A solar array is used as an example of such a system. The emphasis in this review is on the discrepancies that exist in this technology in both flight and ground experiment data. It has been found that, in ground testing, there are facility effects, cell size effects and area scaling uncertainties. For space applications there are area scaling and discharge concerns for an array as well as the influence of the large space structures on the collection process. There are still considerable uncertainties in the high voltage-space plasma interaction technology even after several years of effort.

Stevens, N. John↗
NASA NTRS1989

Carousel deployment mechanism for coilable lattice truss

The development of a mechanism for instrumentation and solar-array deployment is discussed. One part of the technology consists of a smart motor which can operate in either an analog mode to provide high speed and torque, or in the stepper mode to provide accurate positioning. The second technology consists of a coilable lattice mast which is deployed and rotated about its axis with a common drive system. A review of the design and function of the system is presented. Structural and thermal test data are included.

Warden, Robert M.↗
NASA NTRS1989

Beamed laser power in support of near-earth missions

It was found that solar-pumped laser-beamed power is lighter than photovoltaic for power requirements of 150 KWe and above, and is competitive with combined photovoltaic/solar-dynamic over the entire power range investigated. A space station supported by laser-beamed power can be a lower-g facility (reduced drag) than with PV or PV + SD power; has greater freedom of orientation (small receiver moves rather than large arrays or concentrators); and requires less structure (arrays, alpha joints, booms) permitting easier control and fewer vibrational modes. Laser power beaming offers a revolutionary concept for planning designing, and powering large orbiting spacecraft.

Conway, Edmund J.↗
NASA NTRS1990

Latest developments in the Advanced Photovoltaic Solar Array Program

In 1985, the Advanced Photovoltaic Solar Array (APSA) Program was established to demonstrate a producible array system with a specific power greater than 130 W/kg at a 10-kW (BOL) power level. The latest program phase completed fabrication and initial functional testing of a prototype wing representative of a full-scale 5-kW (BOL) wing (except truncated in length to about 1 kW), with weight characteristics that could meet the 130-W/kg (BOL) specific power goal using thin silicon solar cell modules and weight-efficient structural components. The wing configuration and key design details are reviewed, along with results from key component-level and wing-level tests. Projections for future enhancements that may be expected through the use of advanced solar cells and structural components are shown. Performance estimates are given for solar electric propulsion orbital transfer missions through the Van Allen radiation belts. The latest APSA program plans are presented.

Stella, Paul M.↗
NASA NTRS1992

Hypervelocity impact effects on solar cells

One of the space hazards of concern is the problem of natural matter and space debris impacting spacecraft. In addition to mechanical damage, impact velocities greater than 5 km/sec can produce shock induced ionization effects with resultant surface charging and complex chemical interactions. The upper limit of the velocity distribution for these particles is on the order of 70 km/sec. The second source of particulate matter is due to the presence of man and the machinery needed to place satellites in orbit. This 'man made' component of the space debris consists of waste, rocket exhaust, and debris caused by satellite break-up. Most of the particles are small. However as the size increases, debris purposefully thrown overboard such as garbage and human waste, combined with paint chips, plastic, wire fragments, bolts, etc., become formidable hazards which completely dominate the distribution function for some orbits. These larger fragments can produce penetration and spalling of the thick metallic structures associated with spacecraft. The particles most often encountered are aluminum oxide, associated with fuel residue, and paint chips. These debris types can have a wide range of particle sizes. It has been stated that the design of spacecraft will have to take the debris evolution into account and provide additional suitable armor for key components in the near future. The purpose of this work was to subject samples from solar power arrays, one of the key components of any spacecraft, to a debris flux typical of what might be found in space, and measure the degradation of the power panels after impact.

Rose, M. Frank↗