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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 NTRS1999

Pucksat Payload Carrier

There is an ever-expanding need to provide economical space launch opportunities for relatively small science payloads. To address this need, a team at NASA's Goddard Space Flight Center has designed the Pucksat. The Pucksat is a highly versatile payload carrier structure compatible for launching on a Delta II two-stage vehicle as a system co-manifested with a primary payload. It is also compatible for launch on the Air Force Medium Class EELV. Pucksat's basic structural architecture consists of six honeycomb panels attached to six longerons in a hexagonal manner and closed off at the top and bottom with circular rings. Users may configure a co-manifested Pucksat in a number of ways. As examples, co-manifested configurations can be designed to accommodate dedicated missions, multiple experiments, multiple small deployable satellites, or a hybrid of the preceding examples. The Pucksat has fixed lateral dimensions and a downward scaleable height. The dimension across the panel hexagonal flats is 62 in. and the maximum height configuration dimension is 38.5 in. Pucksat has been designed to support a 5000 lbm primary payload, with the center of gravity located no greater than 60 in. from its separation plane, and to accommodate a total co-manifested payload mass of 1275 lbm.

Milam, M. Bruce↗
NASA NTRS1986

Optimization of spherical facets for parabolic solar concentrators

Solar concentrator designs which employ deployable hexagonal panels are being developed for space power systems. An offset optical configuration has been developed which offers significant system level advantages over previously proposed collector designs for space applications. Optical analyses have been performed which show offset reflector intercept factors to be only slightly lower than those for symmetric reflectors with the same slope error. Fluxes on the receiver walls are asymmetric but manageable by varying the tilt angle of the receiver. Greater producibility is achieved by subdividing the hexagonal panels into triangular mirror facets of spherical contour. Optical analysis has been performed upon these to yield near-optimum sizes and radii.

White, J. E.↗
NASA NTRS1981

Summary of LSST systems analysis and integration task for SPS flight test articles

The structural and equipment requirements for two solar power satellite (SPS) test articles are defined. The first SPS concept uses a hexagonal frame structure to stabilize the array of primary tension cables configured to support a Mills Cross antenna containing 17,925 subarrays composed of dipole radiating elements and solid state power amplifier modules. The second test article consists of a microwave antenna and its power source, a 20 by 200 m array of solar cell blankets, both of which are supported by the solar blanket array support structure. The test article structure, a ladder, is comprised of two longitudinal beams (215 m long) spaced 10 m apart and interconnected by six lateral beams. The system control module structure and bridge fitting provide bending and torsional stiffness, and supplement the in plane Vierendeel structure behavior. Mission descriptions, construction, and structure interfaces are addressed.

Greenberg, H. S.↗
NASA NTRS1986

The large deployable reflector: A NASA submillimeter-infrared orbiting observatory

The Large Deployable Reflector (LDR) concept is outlined. The LDR telescope is based on a 20-m diameter reflector. The primary mirror is a filled aperture made up of 84 hexagonal panels, each 2 m edge-to-edge. The panels are based on lightweight structural composite materials. The optical configuration is a four mirror two stage system. The primary mirror is passive. The active optical elements for figure control are at the quaternary mirror. The primary mirror panels are supported by a deployable PAC truss backup structure at the vertices of each hexagon. The four focal plane instruments covering the range of 30 to 1000 microns are located near the vertex of the primary mirror. Some instruments will be cooled with stored cryogens to liquid helium temperatures, others to liquid nitrogen temperatures. The spacecraft functions will be located in a resource module behind the primary mirror. The LDR will be transferred to orbit by the space transportation system and assembled and tested at the space station. It will then be boosted to an orbit of greater than or = 700 km as a free flyer.

Swanson, Paul N.↗
NASA NTRS1960

Ranger Spacecraft

The thermal-control philosophy of the spacecraft currently under development by the Jet Propulsion Laboratory is design by passive means to maintain all components within the tolerances specified by cognizant engineers. Due to the complexity of the configurations, calculations are) of necessity, fairly generalized and final design is based upon tests in an environmental chamber. The Ranger series spacecraft is designed with a basic structure which is common to all models, with additional hardware to suit the individual mission. This basic structure of Rangers A-1 and A-2 is seen as the hexagonal instrument section, the erectable solar panels, the movable antenna, and the omniantenna. The Ranger A-1 and A-2 configuration is for engineering tests and space-exploration, with the scientific instrumentation isolation requirement dictating the spread-out design. The spacecraft stands 12 feet high, weighs 700 to 800 pounds, and has an internal power of 150 watts. Rangers A-3, A-4, and A-5 are designed to rough land a capsule on the moon. For these, a capsule and retrorocket replace the scientific instruments, occupying the space inside the tower structure. The spacecraft must survive many environments. Chronologically they are: 1) Folded configuration inside an aerodynamic shroud on the pad. 2) Thermal flux from shroud aerodynamically heated during boost phase. 3) Coasting up to 30 minutes attached to Agena stage after booster and shroud are separated. 4) Agena stage burning. 5) Coasting and tumbling after separation from Agena until it passes from earth's shadow. 6) Upon reaching sunlight, panels open and begin sun acquisition. 7) Antenna seeks earth after spacecraft locks onto sun. 8) Space phase- "steady state" with vehicle's vertical axis locked on sun, communicating with earth. The philosophy is to design for the sun-acquired mode, making allowances for the transient conditions.

Thostesen, T. O.↗
NASA NTRS1991

The 4-meter lunar engineering telescope

The 16-meter diffraction limited lunar telescope incorporates a primary mirror with 312 one-meter segments; 3 nanometer active optics surface control with laser metrology and hexapod positioners; a space frame structure with one-millimeter stability; and a hexapod mount for pointing. The design data needed to limit risk in this development can be obtained by building a smaller engineering telescope on the moon with all of the features of the 16-meter design. This paper presents a 4.33-meter engineering telescope concept developed by the Summer 1990 Student Program of the NASA/JHU Space Grant Consortium Lunar Telescope Project. The primary mirror, made up of 18 one-meter hexagonal segments, is sized to provide interesting science as well as engineering data. The optics are configured as a Ritchey-Chretien with a coude relay to the focal plane beneath the surface. The optical path is continuously monitored with 3-nanometer precision interferometrically. An active optics processor and piezoelectric actuators operate to maintain the end-to-end optical configuration established by wave front sensing using a guide star. The mirror segments, consisting of a one-centimeter thick faceplate on 30-cm deep ribs, maintain the surface figure to a few nanometers under lunar gravity and thermal environment.

Peacock, Keith↗
NASA NTRS1982

Developing a scalable inert gas ion thruster

Analytical studies to identify and then design a high performance scalable ion thruster operating with either argon or xenon for use in large space systems are presented. The magnetoelectrostatic containment concept is selected for its efficient ion generation capabilities. The iterative nature of the bounding magnetic fields allows the designer to scale both the diameter and length, so that the thruster can be adapted to spacecraft growth over time. Three different thruster assemblies (conical, hexagonal and hemispherical) are evaluated for a 12 cm diameter thruster and performance mapping of the various thruster configurations shows that conical discharge chambers produce the most efficient discharge operation, achieving argon efficiencies of 50-80% mass utilization at 240-310 eV/ion and xenon efficiencies of 60-97% at 240-280 eV/ion. Preliminary testing of the large 30 cm thruster, using argon propellant, indicates a 35% improvement over the 12 cm thruster in mass utilization efficiency. Since initial performance is found to be better than projected, a larger 50 cm thruster is already in the development stage.

James, E.↗
NASA NTRS1998

Static Thrust and Vectoring Performance of a Spherical Convergent Flap Nozzle with a Nonrectangular Divergent Duct

The static internal performance of a multiaxis-thrust-vectoring, spherical convergent flap (SCF) nozzle with a non-rectangular divergent duct was obtained in the model preparation area of the Langley 16-Foot Transonic Tunnel. Duct cross sections of hexagonal and bowtie shapes were tested. Additional geometric parameters included throat area (power setting), pitch flap deflection angle, and yaw gimbal angle. Nozzle pressure ratio was varied from 2 to 12 for dry power configurations and from 2 to 6 for afterburning power configurations. Approximately a 1-percent loss in thrust efficiency from SCF nozzles with a rectangular divergent duct was incurred as a result of internal oblique shocks in the flow field. The internal oblique shocks were the result of cross flow generated by the vee-shaped geometric throat. The hexagonal and bowtie nozzles had mirror-imaged flow fields and therefore similar thrust performance. Thrust vectoring was not hampered by the three-dimensional internal geometry of the nozzles. Flow visualization indicates pitch thrust-vector angles larger than 10' may be achievable with minimal adverse effect on or a possible gain in resultant thrust efficiency as compared with the performance at a pitch thrust-vector angle of 10 deg.

Wing, David J.↗
NASA NTRS1974

Some injector element detail effects on screech in hydrogen-oxygen rockets

An experimental investigation was conducted to learn more about how the specific details of a concentric tube injection element affect the screech characteristics of a hydrogen-oxygen rocket engine. The four variables investigated were (1) impingement angle, (2) oxidizer tube blunt base thickness, (3) oxidizer tube recess and extension, and (4) oxidizer tube-annulus concentricity. Tests were made using a 27.34-cm (10.77-in.) diameter heat sink combustor at nominally 300-psia chamber pressure. All of the test variables were investigated using a 157 element circular pattern injector. Additional oxidizer tube recess tests were made with a 421 element hexagonal pattern injector. Tests were conducted over the oxidant-fuel ratio range of 4 to 6. Stability evaluation for each configuration was made using the hydrogen temperature rating technique. Several element configurations were also cold flow tested using nitrogen and water as simulants. The element detail changes resulted in changes in hydrogen injector pressure drop even though the physical injection area was constant for all similar tests in both hot firing and cold flow tests. These changes in injector pressure drop produced changes in combustion stability. The data were correlated with a modified version of a previously reported injection area ratio correlation. By interpreting changes in injector pressure drop as changes in injector hydrogen flow resistance, the data were compared with a hydrogen flow response stability model and were found to be in agreement.

Hannum, N. P.↗
NASA NTRS1979

Concept for UF6-fueled self-critical DNPL reactor system

An analytical study of a self-critical nuclear pumped laser system concept was performed. The primary emphasis was on a reactor concept which would use gaseous uranium hexafluoride (UF6) as the fissioning material. A reference configuration was selected which has a 3.2 cu m lasing volume as the reactor core. The core is composed of a series of hexagonal graphite tubes which are surrounded by a reflector-moderator composed either of heavy water or beryllium. Laser transitions requiring average fission power densities less than approximately 1 kW/cu cm for excitation are most attractive. Operation at wavelengths greater than approximately 400 nm may be required because of limitations imposed by the opacity of gaseous UF6. Further research directed toward identification of UF6 compatible lasing transitions is required.

Rodgers, R. J.↗
NASA NTRS1999

Mars Orbiter Sample Return Power Design

The NASA/JPL 2003/2005 Mars Sample Return (MSR) Missions will each have a sample return canister that will be filled with samples cored from the surface of MARS. These spherical canisters will be 14.8 cm in diameter and must be powered only by solar cells on the surface and must communicate using RF transmission with the recovery vehicle that will be coming in 2006 or 2009 to retrieve the canister. This paper considers the aspect and conclusion that went into the design of the power system that achieves the maximum power with the minimum risk. The power output for the spherical orbiting canister was modeled and plotted in various views of the orbit by the SOAP program developed by JPL. The requirements and geometry for a solar array on a sphere are unique and place special constraints on the design. These requirements include 1) accommodating a lid for sample loading into the canister, surface area was restricted from use on the Northern pole of the spherical canister. 2) minimal cell surface coverage (maximum cell efficiency), less than 40%, for recovery vehicle to locate the canister by optical techniques. 3) a RF transmission during 50% of MARS orbit time on any spin axis, which requires optimum circuit placement of the solar cell onto the spherical canister. The best configuration would have been a 4.5 volt round cell, but in the real world we compromised with six triangular silicon cells connected in series to form a hexagon. These hexagon circuits would be mounted onto a flat facet cut into the spherical canister. The surface flats are required in order to maximize power, the surface of the cells connected in series must be at the same angle relative to the sun. The flat facets intersect each other to allow twelve circuits evenly spaced just North and twelve circuits South of the equator of the spherical canister. Connecting these circuits in parallel allows sufficient power to operate the transmitter at minimum solar exposure, Northern pole of the canister facing the sun. Additional power, as much as 20%, is also generated by the circuits facing MARS due to albedo of MARS.

Mardesich, N.↗
NASA NTRS1980

Studies of H2O on beta-AgI surfaces - An effective pair potential model

The adsorption of a water molecule on surfaces of beta-AgI, the hexagonal crystal believed to be primarily responsible for the ice-nucleating properties of AgI, is studied on the basis of an effective pair potential model. The water molecule is represented by a rigid point charge ST-2 model and the AgI substrate by an array of point atoms, and maximal binding energy surfaces and optimal H2O configurations are generated for the water molecule adsorbed on the rigid and unrelaxed basal and prism faces. Modeling of H2O adsorption above a two-layer ledge, an iodine vacancy and an H2O molecule trapped in the vacancy indicates that H2O adsorption is favored at interstitial sites where no substrate atoms lie directly below. The prism face is found to attract the water molecule more strongly and provide larger energy barriers to surface diffusion than basal face sites, with the ideal basal faces providing hexagonal patterns of adsorption sites for the H2O with preferred dipole moments aligned.

Hale, B. N.↗
NASA NTRS1972

Heat transfer to four fineness-ratio-1.6 hexagonal prisms with various corner radii at Mach 6

An investigation was conducted in the Langley 20-inch Mach 6 tunnel to define the aerodynamic heat transfer to the radioisotope fuel cask (heat source) of the SNAP-19/Pioneer power system. The shape of the SNAP-19/Pioneer heat source is that of a hexagonal prism with flat ends; the fineness ratio, based on maximum (edge to edge) diameter, is 1.61. Phase-change-paint heat-transfer data and schlieren photographs were obtained on four possible 1/2-scale entry configurations of the SNAP-19/Pioneer heat source. Tests were conducted over a wide range of attitudes and at nominal Reynolds numbers, based on the length of the unablated configuration, of 33,000; 84,000; and 2,200,000.

Hunt, J. L.↗
NASA NTRS2004

Multifunctional Inflatable Structure Being Developed for the PowerSphere Concept

NASA has funded a collaborative team of The Aerospace Corporation, ILC Dover, Lockheed Martin, and NASA Glenn Research Center to develop the Multifunctional Inflatable Structure (MIS) for a "PowerSphere" concept through a NASA Research Announcement. This power system concept has several advantages, including a high collection area, low weight and stowage volume, and the elimination of all solar array pointing mechanisms. The current 3-year effort will culminate with the fabrication and testing of a fully functional engineering development unit. The baseline design of the Power-Sphere consists of two opposing semispherical domes connected to a central spacecraft. Each semispherical dome consists of hexagonal and pentagonal solar cell panels that together form a geodetic sphere. Inflatable ultraviolet (UV) rigidizable tubular hinges between the solar cell panels and UV rigidizable isogrid center columns with imbedded flex circuitry form the MIS. The reference configuration for the PowerSphere is a 0.6-m-diameter (fully deployed) spacecraft with a total mass budget of 4 kg (1 kg for PowerSphere, 3 kg for spacecraft) capable of producing 29 W of electricity with 10-percent-efficient thin-film solar cells. In a stowed configuration, the solar cell panels will be folded sequentially to the outside of the instrument decks. The center column will be z-folded between the instrument decks and the spacecraft housing for packaging. The instrument panel will secure the z-folded stack with launch ties. After launch, once the release tie is triggered, the center column and hinge tubes will inflate and be rigidized in their final configurations by ultraviolet radiation. The overall PowerSphere deployment sequence is shown pictorially in the following illustration.

Peterson, Todd T.↗
NASA NTRS2009

Connecting Node

A paper describes the Octanode, a connecting node that facilitates the integration of multiple docking mechanisms, hatches, windows, and internal and external systems with the use of flat surfaces. The Octanode is a 26- faced Great Rhombicuboctahedron Archi medean solid with six octagonshaped panels, eight hexagon-shaped panels, and 12 square panels using three unique, simple, flat shapes to construct a spherical approximation. Each flat shape can be constructed with a variety of material and manufacturing techniques, such as honeycomb composite panels or a pocketed skinstringer configuration, using conventional means. The flat shapes can be connected together and sealed to create a pressurizable volume by the use of any conventional means including welding or fastening devices and sealant. The node can then be connected to other elements to allow transfer between those elements, or it could serve as an airlock. The Octanode can be manufactured on the ground and can be integrated with subsystems including hatches and ports. The node can then be transported to its intended location, whether on orbit or on surface. Any of the flat panels could be replaced by curved ones, turning the node into a copula. Windows may be placed on flat panes with optimal viewing angles that are not blocked by large connecting nodes. The advantage of using flat panels to represent a spherical approximation is that this allows for easier integration of subsystems and design features.

Johnson, Christopher J.↗
NASA NTRS1991

Molecular dynamics and quasidynamics simulations of the annealing of bulk and near-surface interstitials formed in molecular-beam epitaxial Si due to low-energy particle bombardment during deposition

The relaxation, diffusion, and annihilation of split and hexagonal interstitials resulting from 10 eV Si irradiation of (2x1)-terminated Si(100) are investigated. Molecular dynamics and quasidynamics simulations, utilizing the Tersoff many-body potential are used in the investigation. The interstitials are created in layers two through six, and stable atomic configurations and total potential energies are derived as a function of site symmetry and layer depth. The interstitial Si atoms are allowed to diffuse, and the total potential energy changes are calculated. Lattice configurations along each path, as well as the starting configurations, are relaxed, and minimum energy diffusion paths are derived. The results show that the minimum energy paths are toward the surface and generally involved tetrahedral sites. The calculated interstitial migration activation energies are always less than 1.4 eV and are much lower in the near-surface region than in the bulk.

Kitabatake, M.↗