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At least 469 records · Page 26

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NASA NTRS2007

Hypervelocity Impact (HVI): WLE High Fidelity Specimen Fg(RCC)-2 - Volume 6

During 2003 and 2004, the Johnson Space Center's White Sands Testing Facility in Las Cruces, New Mexico conducted hypervelocity impact tests on the space shuttle wing leading edge. Hypervelocity impact tests were conducted to determine if Micro-Meteoroid/Orbital Debris impacts could be reliably detected and located using simple passive ultrasonic methods. The objective of Target Fg(RCC)-2 was to study hypervelocity impacts through the reinforced carbon-carbon (RCC) panels of the Wing Leading Edge. Fiberglass was used in place of RCC in the initial tests. Impact damage was detected using lightweight, low power instrumentation capable of being used in flight.

Gorman, Michael R.↗

NASA NTRS1981

Temperature stabilized linkage

A temperature compensated passive linkage for interconnecting two members having different coefficients of expansion. Preferably the linkage is utilized for interconnecting the stabilizer of a plane with its elevator. The stabilizer may be a graphite epoxy composite while the elevator may consist of aluminum-fiberglass. The differences in the rate of expansion of the two members is compensated by the linkage of the invention which in turn will move a side load hinge fitting to minimize aerodynamic and mechanical problems.

Ohgi, George Y.↗

NASA NTRS1985

Unitary plate electrode

The unitary electrode (10) comprises a porous sheet (12) of fiberglass the strands (14) of which contain a coating (16) of conductive tin oxide. The lower portion of the sheet contains a layer (18) of resin and the upper layer (20) contains lead dioxide forming a positive active electrode on an electrolyte-impervious layer. The strands (14) form a continuous conduction path through both layers (16, 18). Tin oxide is prevented from reduction by coating the surface of the plate facing the negative electrode with a conductive, impervious layer resistant to reduction such as a thin film (130) of lead or graphite filled resin adhered to the plate with a layer (31) of conductive adhesive. The plate (10) can be formed by casting a molten resin from kettle (60) onto a sheet of glass wool (56) overlying a sheet of lead foil and then applying positive active paste from hopper (64) into the upper layer (68). The plate can also be formed by passing an assembly of a sheet ( 80) of resin, a sheet (86) of sintered glass and a sheet (90) of lead between the nip (92) of heated rollers (93, 95) and then filling lead oxide into the pores (116) of the upper layer (118).

Rowlette, John J.↗

NASA NTRS1985

Sealed bipolar multi-cell battery

A low maintenance battery comprises a sealed casing (102) having a cavity (204) receiving a stack (108) of monopolar and bipolar plates (110, 112) interspersed with mats (122) of fiberglass felt immobilizing electrolyte and forming a gas path for diffusion of oxygen to the negative electrode (112). The casing contains a single, resealable vent (202). During charging, oxygen generated at the positive plate (110) diffuses through the mat (122) and decomposes at the negative plate (112). The negative plate (112) has a larger capacity than the positive plate (110) to prevent evolution of hydrogen. Voltage potential is increased or decreased by placing one or more bipolar plates (120) in series relation between alternating positive and negative monopolar plates that are connected in parallel to respective positive and negative battery terminals (104, 106). The monoplates (110, 112) are connected in parallel to respective positive and negative battery terminals (104, 106) by way of bus plates (114, 116, 130, 132). The resealable vent permits operation by way of the oxygen cycle which prevents loss of electrolyte. The vent prevents invasion of oxygen from the ambient to discharge negative plates.

Rowlette, John J.↗

NASA NTRS1985

Battery plate containing filler with conductive coating

The plate (10) comprises a matrix or binder resin phase (12) in which is dispersed particulate, conductive tin oxide such as tin oxide coated glass fibers (14). A monopolar plate (11) is prepared by coating a layer (18) of electrolytically active material onto a surface of the plate (10). Tin oxide is prevented from reduction by coating a surface of the plate (10) with a conductive, impervious layer resistant to reduction such as a thin film (22) of lead adhered to the plate with a layer (21) of conductive adhesive. The plate (10) can be formed by casting a molten dispersion from mixer (36) onto a sheet (30) of lead foil or by passing an assembly of a sheet (41) of resin, a sheet (43) of fiberglass and a sheet (45) of lead between the nip of heated rollers (48, 50).

Rowlette, John J.↗

NASA NTRS1976

Method of manufacturing an overwrapped pressure vessel

A pressure vessel of the type wherein a metallic liner in the shape of a cylindrical portion with a dome-shaped portion at each end thereof is overwrapped by a plurality of layers of resin coated, single fiberglass filaments. A four-step wrapping technique reinforces the vessel with overwrap material at the most likely areas for vessel failure. Overwrapping of the vessel is followed by a sizing pressurization cycle which induces a compressive prestress into the liner and thereby permits the liner to deform elastically through an increased strain range.

Beck, Emory J.↗

NASA NTRS1986

Battery plate containing filler with conductive coating

Rowlette, John J.↗

NASA NTRS1989

Positive paste with lead-coated glass fibers

A reliable positive plate for deep cycle lead-acid batteries is provided by dispersing about 2 percent by volume of fine, short, lead-coated reinforcement fibers such as fiberglass in the positive active paste before being applied to the grid of a battery plate and cured. The fibers improve the structural integrity of the plate and also provide electrical paths in a discharged plate. This results in greater utilization of the lead dioxide active material by preventing charged portions of the plate from becoming electrically isolated.

Rippel, Wally E.↗

NASA NTRS2006

Viking Afterbody Heating Computations and Comparisons to Flight Data

Computational fluid dynamics predictions of Viking Lander 1 entry vehicle afterbody heating are compared to flight data. The analysis includes a derivation of heat flux from temperature data at two base cover locations, as well as a discussion of available reconstructed entry trajectories. Based on the raw temperature-time history data, convective heat flux is derived to be 0.63-1.10 W/cm2 for the aluminum base cover at the time of thermocouple failure. Peak heat flux at the fiberglass base cover thermocouple is estimated to be 0.54-0.76 W/cm2, occurring 16 seconds after peak stagnation point heat flux. Navier-Stokes computational solutions are obtained with two separate codes using an 8- species Mars gas model in chemical and thermal non-equilibrium. Flowfield solutions using local time-stepping did not result in converged heating at either thermocouple location. A global time-stepping approach improved the computational stability, but steady state heat flux was not reached for either base cover location. Both thermocouple locations lie within a separated flow region of the base cover that is likely unsteady. Heat flux computations averaged over the solution history are generally below the flight data and do not vary smoothly over time for both base cover locations. Possible reasons for the mismatch between flight data and flowfield solutions include underestimated conduction effects and limitations of the computational methods.

Edquist, Karl T.↗

NASA NTRS2008

Active Co-Storage of Cryogenic Propellants for Lunar Explortation

Long-term storage of cryogenic propellants is a critical requirement for NASA's effort to return to the moon. Liquid hydrogen and liquid oxygen provide the highest specific impulse of any practical chemical propulsion system, and thus provides the greatest payload mass per unit of launch mass. Future manned missions will require vehicles with the flexibility to remain in orbit for months, necessitating long-term storage of these cryogenic liquids. For decades cryogenic scientific satellites have used dual cryogens with different temperatures to cool instruments. This technology utilizes a higher temperature cryogen to provide a stage that efficiently intercepts a large fraction of the heat that would otherwise be incident on the lower temperature cryogen. This interception reduces the boil-off of the lower temperature cryogen and increasing the overall life-time of the mission. The Active Co-Storage concept is implemented similarly; the 101 K liquid oxygen thermally shields the 24 K liquid hydrogen. A thermal radiation shield that is linked to the liquid oxygen tank shrouds the liquid hydrogen tank, thereby preventing the liquid hydrogen tank from being directly exposed to the 300 K external environment. Modern cryocooler technology can eliminate the liquid oxygen boil-off and also cool the thermal radiation shield thereby reducing the liquid hydrogen boil-off to a small fraction of the unshielded rate. The thermal radiation shield can be a simple conductive shroud or a more sophisticated but lighter Broad Area Cooling (BAC) shroud. The paper describes the design impact of an active co-storage system for the Altair Descent Vehicle. This paper also compares the spacecraft-level impacts of the conductive shroud and the BAC shroud active co-storage concepts with a passive storage option in the context of the different scales of spacecraft that will be used for the lunar exploration effort - the Altair Ascent and Descent Vehicles, the Orion, and the Ares V Earth Departure Stage. The paper also reports on a subscale test of this active co-storage configuration. The test tank is 0.7 m in diameter, approximately one-third the dimension of tanks that would be needed in a lunar ascent module. A thin-walled fiberglass skirt supports and isolates the tank from a 100 K stage. A similar thin-walled skirt supports the lOOK stage from the ambient temperature structure. An aluminum shield with a heavy MLI blanket surrounds the tank and is attached at the 100 K stage. In this initial phase of the project, there is no tank on the 100 K stage, but it is actively cooled by a single-stage cryocooler similar in design to the one used on the RHESSI mission. The test configuration includes a number of innovative elements, including a helical support heat exchanger and an external thermodynamic vent/heat interception system. To avoid the complexity of an explosive gas handling system, testing will be done with liquid helium and liquid neon as simulant fluids. The properties of these fluids bracket the properties of liquid hydrogen. Instrumentation allows tank temperature and shield temperature profiles, tank liquid levels, and pressure drops through the flow lines, to be measured.

Mustafi, S.↗

NASA NTRS2008

Microwave Nondestructive Evaluation of Dielectric Materials with a Metamaterial Lens

A novel microwave Nondestructive Evaluation (NDE) sensor was developed in an attempt to increase the sensitivity of the microwave NDE method for detection of defects small relative to a wavelength. The sensor was designed on the basis of a negative index material (NIM) lens. Characterization of the lens was performed to determine its resonant frequency, index of refraction, focus spot size, and optimal focusing length (for proper sample location). A sub-wavelength spot size (3 dB) of 0.48 lambda was obtained. The proof of concept for the sensor was achieved when a fiberglass sample with a 3 mm diameter through hole (perpendicular to the propagation direction of the wave) was tested. The hole was successfully detected with an 8.2 cm wavelength electromagnetic wave. This method is able to detect a defect that is 0.037 lambda. This method has certain advantages over other far field and near field microwave NDE methods currently in use.

Shreiber, Daniel↗

NASA NTRS2009

Tensile Fabrics Enhance Architecture Around the World

Using a remarkable fabric originally developed to protect Apollo astronauts, Birdair Inc. of Amherst, New York, has crafted highly durable, safe, environmentally friendly, and architecturally stunning tensile membrane roofs for over 900 landmark structures around the world. Travelers in airports, sports fans at stadiums, and shoppers in malls have all experienced the benefits of the Teflon-coated fiberglass fabric that has enabled Birdair to grow from a small company established in its founder?s kitchen in 1955 to a multimillion-dollar specialty contractor today.

Source record↗

NASA NTRS2007

Composite Pulse Tube

A modification of the design of the pulse tube in a pulse-tube cryocooler reduces axial thermal conductance while preserving radial thermal conductance. It is desirable to minimize axial thermal conductance in the pulse-tube wall to minimize leakage of heat between the warm and cold ends of the pulse tube. At the same time, it is desirable to maximize radial thermal conductance at the cold end of the pulse tube to ensure adequate thermal contact between (1) a heat exchanger in the form of a stack of copper screens inside the pulse tube at the cold end and (2) the remainder of the cold tip, which is the object to which the heat load is applied and from which heat must be removed. The modified design yields a low-heat-leak pulse tube that can be easily integrated with a cold tip. A typical pulse tube of prior design is either a thin-walled metal tube or a metal tube with a nonmetallic lining. It is desirable that the outer surface of a pulse tube be cylindrical (in contradistinction to tapered) to simplify the design of a regenerator that is also part of the cryocooler. Under some conditions, it is desirable to taper the inner surface of the pulse tube to reduce acoustic streaming. The combination of a cylindrical outer surface and a tapered inner surface can lead to unacceptably large axial conduction if the pulse tube is made entirely of metal. Making the pulse-tube wall of a nonmetallic, lowthermal- conductivity material would not solve the problem because the wall would not afford the needed thermal contact for the stack of screens in the cold end. The modified design calls for fabricating the pulse tube in two parts: a longer, nonmetallic part that is tapered on the inside and cylindrical on the outside and a shorter, metallic part that is cylindrical on both the inside and the outside. The nonmetallic part can be made from G-10 fiberglass-reinforced epoxy or other low-thermal-conductivity, cryogenically compatible material. The metallic part must have high thermal conductivity in the cryogenic temperature range and would typically be made of pure copper to satisfy this requirement. The metallic part is bonded to the nonmetallic part with epoxy. Copper screens are inserted in the metallic part to form the cold-end heat exchanger, then the assembled pulse tube is inserted in the cold tip.

Martin, Jerry L.↗

NASA NTRS2010

Design and Analysis of Outer Mold Line Close-outs for the Max Launch Abort System (MLAS) Flight Experiment

In 2007, the NASA Exploration Systems Mission Directorate (ESMD) chartered the NASA Engineering Safety Center (NESC) to demonstrate an alternate launch abort concept as risk mitigation for the Orion project's baseline "tower" design. On July 8, 2009, a full scale, passive aerodynamically stabilized Max Launch Abort System (MLAS) pad abort demonstrator was successfully launched from NASA Goddard Space Flight Center's Wallops Flight Facility. Aerodynamic close-outs were required to cover openings on the MLAS fairing to prevent aerodynamic flow-through and to maintain the MLAS OML surface shape. Two-ply duct tape covers were designed to meet these needs. The duct tape used was a high strength fiber reinforced duct tape with a rubberized adhesive that demonstrated 4.6 lb/in adhesion strength to the unpainted fiberglass fairing. Adhesion strength was observed to increase as a function of time. The covers were analyzed and experimentally tested to demonstrate their ability to maintain integrity under anticipated vehicle ascent pressure loads and to not impede firing of the drogue chute mortars. Testing included vacuum testing and a mortar fire test. Tape covers were layed-up on thin Teflon sheets to facilitate installation on the vehicle. Custom cut foam insulation board was used to fill mortar hole and separation joint cavities and provide support to the applied tape covers. Flight test results showed that the tape covers remained adhered during flight.

Woods-Vedeler, Jessica A.↗

NASA NTRS2010

Evaluation of the Water Film Weber Number in Glaze Icing Scaling

Icing scaling tests were performed in the NASA Glenn Icing Research Tunnel to evaluate a new scaling method, developed and proposed by Feo for glaze icing, in which the scale liquid water content and velocity were found by matching reference and scale values of the nondimensional water-film thickness expression and the film Weber number. For comparison purpose, tests were also conducted using the constant We(sub L) method for velocity scaling. The reference tests used a full-span, fiberglass, 91.4-cm-chord NACA 0012 model with velocities of 76 and 100 knot and MVD sizes of 150 and 195 microns. Scale-to-reference model size ratio was 1:2.6. All tests were made at 0deg AOA. Results will be presented for stagnation point freezing fractions of 0.3 and 0.5.

Tsao, Jen-Ching↗

NASA NTRS2010

Steerable Hopping Six-Legged Robot

The figure depicts selected aspects of a six-legged robot that moves by hopping and that can be steered in the sense that it can be launched into a hop in a controllable direction. This is a prototype of hopping robots being developed for use in scientific exploration of rough terrain on remote planets that have surface gravitation less than that of Earth. Hopping robots could also be used on Earth, albeit at diminished hopping distances associated with the greater Earth gravitation. The upper end of each leg is connected through two universal joints to an upper and a lower hexagonal frame, such that the tilt of the leg depends on the relative position of the two frames. Two non-back-driveable worm-gear motor drives are used to control the relative position of the two frames along two axes 120 apart, thereby controlling the common tilt of all six legs and thereby, further, controlling the direction of hopping. Each leg includes an upper and a lower aluminum frame segment with a joint between them. A fiberglass spring, connected via hinges to both segments, is used to store hopping energy prior to launch into a hop and to cushion the landing at the end of the hop. A cable for loading the spring is run into each leg through the center of the universal joints and then down along the center lines of the segments to the lower end of the leg. A central spool actuated by a motor with a harmonic drive and an electromagnetic clutch winds in all six cables to compress all six springs (thereby also flexing all six legs) simultaneously. To ensure that all the legs push off and land in the same direction, timing- belt pulley drives are attached to the leg segments, restricting the flexing and extension of all six legs to a common linear motion. In preparation for a hop, the spool can be driven to load the spring legs by an amount corresponding to a desired hop distance within range. The amount of compression can be computed from the reading of a shaft-angle encoder that indicates the amount by which the spool has been turned. When the robot is ready to hop, the electromagnetic clutch disengages the motor from the spool, thus releasing the cable restraints on the springs and allowing the springs to extend all six legs simultaneously.

Younse, Paulo↗

NASA NTRS2010

Cryogenic Vacuum Insulation for Vessels and Piping

Cryogenic vacuum insulation systems, with proper materials selection and execution, can offer the highest levels of thermal performance. Three areas of consideration are vital to achieve the optimum result: materials, representative test conditions, and engineering approach for the particular application. Deficiency in one of these three areas can prevent optimum performance and lead to severe inefficiency. Materials of interest include micro-fiberglass, multilayer insulation, and composite arrangements. Cylindrical liquid nitrogen boil-off calorimetry methods were used. The need for standard thermal conductivity data is addressed through baseline testing. Engineering analysis and design factors such as layer thickness, density, and practicality are also considered.

Kogan, A.↗

NASA NTRS2011

Helium-Cooled Black Shroud for Subscale Cryogenic Testing

This shroud provides a deep-space simulating environment for testing scaled-down models of passively cooling systems for spaceflight optics and instruments. It is used inside a liquid-nitrogen- cooled vacuum chamber, and it is cooled by liquid helium to 5 K. It has an inside geometry of approximately 1.6 m diameter by 0.45 m tall. The inside surfaces of its top and sidewalls have a thermal absorptivity greater than 0.96. The bottom wall has a large central opening that is easily customized to allow a specific test item to extend through it. This enables testing of scale models of realistic passive cooling configurations that feature a very large temperature drop between the deepspace-facing cooled side and the Sun/Earth-facing warm side. This shroud has an innovative thermal closeout of the bottom wall, so that a test sample can have a hot (room temperature) side outside of the shroud, and a cold side inside the shroud. The combination of this closeout and the very black walls keeps radiated heat from the sample s warm end from entering the shroud, reflecting off the walls and heating the sample s cold end. The shroud includes 12 vertical rectangular sheet-copper side panels that are oriented in a circular pattern. Using tabs bent off from their edges, these side panels are bolted to each other and to a steel support ring on which they rest. The removable shroud top is a large copper sheet that rests on, and is bolted to, the support ring when the shroud is closed. The support ring stands on four fiberglass tube legs, which isolate it thermally from the vacuum chamber bottom. The insides of the cooper top and side panels are completely covered with 25- mm-thick aluminum honeycomb panels. This honeycomb is painted black before it is epoxied to the copper surfaces. A spiral-shaped copper tube, clamped at many different locations to the outside of the top copper plate, serves as part of the liquid helium cooling loop. Another copper tube, plumbed in a series to the top plate s tube, is clamped to the sidewall tabs where they are bolted to the support ring. Flowing liquid helium through these tubes cools the entire shroud to 5 K. The entire shroud is wrapped loosely in a layer of double-aluminized Kapton. The support ring s inner diameter is the largest possible hole through which the test item can extend into the shroud. Twelve custom-sized trapezoidal copper sheets extend inward from the support ring to within a few millimeters of the test item. Attached to the inner edge of each of these sheets is a custom-shaped strip of Kapton, which is aluminum- coated on the warm-facing (outer) side, and has thin Dacron netting attached to its cold-facing side. This Kapton rests against the test item, but the Dacron keeps it from making significant thermal contact. The result is a non-contact, radiatively reflective thermal closeout with essentially no gap through which radiation can pass. In this way, the part of the test item outside the shroud can be heated to relatively high temperatures without any radiative heat leaking to the inside.

Tuttle, James↗