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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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At least 343 records · Page 19
All sourcesNASA NTRSDOE OSTI
NASA NTRS1989

Solar dynamic power for Space Station Freedom

The Space Station Freedom Program is presently planned to consist of two phases. At the completion of Phase 1, Freedom's manned base will consist of a transverse boom with attached manned modules and 75 kW of available electric power supplied by photovoltaic (PV) power sources. In Phase 2, electric power available to the manned base will be increased to 125 kW by the addition of two solar dynamic (SD) power modules, one at each end of the transverse boom. Power for manned base growth beyond Phase 2 will be supplied by additional SD modules. Studies show that SD power for the growth eras will result in life cycle cost savings of $3 to $4 billion when compared to PV-supplied power. In the SD power modules for Space Station Freedom, an offset parabolic concentrator collects and focuses solar energy into a heat receiver. To allow full power operation over the entire orbit, the receiver includes integral thermal energy storage by means of the heat of fusion of a salt mixture. Thermal energy is removed from the receiver and converted to electrical energy by a power conversion unit (PCU) which includes a closed brayton cycle (CBC) heat engine and an alternator. The receiver/PCU/radiator combination will be completely assembled and charged with gas and cooling fluid on earth before launch to orbit. The concentrator subassemblies will be pre-aligned and stowed in the orbiter bay before launch. On orbit, the receiver/PCU/radiator assembly will be installed as a unit. The pre-aligned concentrator panels will then be latched together and the total concentrator attached to the receiver/PCU/radiator by the astronauts. After final electric connections are made and checkout is complete, the SD power module will be ready for operation.

Labus, Thomas L.↗
NASA NTRS1989

Large space antenna concepts for ESGP

It is appropriate to note that 1988 marks the 100th anniversary of the birth of the reflector antenna. It was in 1888 that Heinrich Hertz constructed the first one, a parabolic cylinder made of sheet zinc bent to shape and supported by a wooden frame. Hertz demonstrated the existence of the electromagnetic waves that had been predicted theoretically by James Clerk Maxwell some 22 years earlier. In the 100 years since Hertz's pioneering work the field of electromagnetics has grown explosively: one of the technologies is that of remote sensing of planet Earth by means of electromagnetic waves, using both passive and active sensors located on an Earth Science Geostationary Platform (ESEP). For these purposes some exquisitely sensitive instruments were developed, capable of reaching to the fringes of the known universe, and relying on large reflector antennas to collect the minute signals and direct them to appropriate receiving devices. These antennas are electrically large, with diameters of 3000 to 10,000 wavelengths and with gains approaching 80 to 90 dB. Some of the reflector antennas proposed for ESGP are also electrically large. For example, at 220 GHz a 4-meter reflector is nearly 3000 wavelengths in diameter, and is electrically quite comparable with a number of the millimeter wave radiotelescopes that are being built around the world. Its surface must meet stringent requirements on rms smoothness, and ability to resist deformation. Here, however, the environmental forces at work are different. There are no varying forces due to wind and gravity, but inertial forces due to mechanical scanning must be reckoned with. With this form of beam scanning, minimizing momentum transfer to the space platform is a problem that demands an answer. Finally, reflector surface distortion due to thermal gradients caused by the solar flux probably represents the most challenging problem to be solved if these Large Space Antennas are to achieve the gain and resolution required of them.

Love, Allan W.↗
NASA NTRS1989

Photovoltaic module on-orbit assembly for Space Station Freedom

One of the elements of the Space Station Freedom power system is the photovoltaic (PV) module. These modules will be assembled on-orbit during the assembly phase of the program. These modules will be assembled either from the Shuttle Orbiter or from the Mobile Servicing Center (MSC). The different types of assembly operations that will be used to assemble PV modules are described.

Sours, T.↗
NASA NTRS1989

Solar dynamic power module design

Studies have shown that the use of solar dynamic (SD) power for the growth areas of the Space Station Freedom program will result in life cycle cost savings when compared to power supplied by photovoltaic sources. In the SD power module, a concentrator collects and focuses solar energy into a heat receiver which has integral thermal energy storage. A Power Conversion Unit (PCU) based on the closed Brayton cycle removes thermal energy from the receiver and converts that energy to electrical energy. Since the closed Brayton cycle is a single phase gas cycle, the conversion hardware (heat exchangers, turbine, compressor, etc.) can be designed for operation in low earth orbit, and tested with confidence in test facilities on earth before launch into space. The concentrator subassemblies will be aligned and the receiver/PCU/radiator combination completely assembled and charged with gas and cooling liquid on earth before launch to, and assembly on, orbit.

Secunde, Richard R.↗
NASA NTRS1991

Heating and thermal transformation of micrometeoroids entering the earth's atmosphere

The present numerical solutions for the atmospheric entry of 10 micron-1 mm diameter micrometeoroids gave attention to ablative mass loss and cooling, together with gravitational and curvature effects, for entry velocities in the 11.2-72 km/sec range. Maximum temperature and mass-loss rates are found to generally occur at altitudes between 85 and 90 km, during about 1 sec of peak heating; the survival of all particles in the 70 micron-1 mm size range is noted to be limited to those with minimal entry velocity. Virtually all of the 'cosmic spherules' of more than 70-mm diameter, as well as giant unmelted micrometeorites, are implied by the present results to be of asteroidal origin.

Love, S. G.↗
NASA NTRS1992

Plotting Program For Aerodynamical Computations

Panel Method Aerodynamic Plotting Program (PMAPP) written to plot results of aerodynamical analyses (flow data) from PMARC (ARC-12642), program for computation of three-dimensional flow by low-order panel method. PMAPP is interactive, color-capable graphics program designed to work with variety of computer terminals and printers. Also used to plot data from other programs, provided files containing those data defined according to PMARC conventions. Written in FORTRAN 77.

Wigren, L. J.↗
NASA NTRS1991

Emission lines in the long period Cepheid l Carinae

For the Cepheid (l) Carinae with a pulsation period of 35.5 days we have studied the emission line fluxes as a function of pulsational phase in order to find out whether we see chromosphere and transition layer emission or whether we see emission due to an outward moving shock. All emission lines show a steep increase in flux shortly before maximum light suggestive of a shock moving through the surface layers. The large ratio of the C IV to C II line fluxes shows that these are not transition layer lines. During maximum light the large ratio of the C IV to C II line fluxes also suggests that we see emission from a shock with velocities greater than 100 km/sec such that C IV emission can be excited. With such velocities mass outflow appears possible. The variations seen in the Mg II line profiles show that there is an internal absorption over a broad velocity band independent of the pulsational phase. We attribute this absorption to a circumstellar 'shell'. This 'shell' appears to be seen also as spatially extended emission in the O I line at 1300 angstrom, which is probably excited by resonance with Ly beta.

Boehm-Vitense, Erika↗
NASA NTRS1992

Changes in ice cover thickness and lake level of Lake Hoare, Antarctica - Implications for local climatic change

Results are reported from 10 years of ice-thickness measurements at perennially ice-covered Lake Hoare in southern Victoria Land, Antarctica. The ice cover of this lake had been thinning steadily at a rate exceeding 20 cm/yr during the last decade but seems to have recently stabilized at a thickness of 3.3 m. Data concerning lake level and degree-days above freezing are presented to show the relationship between peak summer temperatures and the volume of glacier-derived meltwater entering Lake Hoare each summer. From these latter data it is inferred that peak summer temperatures have been above 0 C for a progressively longer period of time each year since 1972. Possible explanations for the thinning of the lake ice are considered. The thickness of the ice cover is determined by the balance between freezing during the winter and ablation that occurs all year but maximizes in summer. It is suggested that the term most likely responsible for the change in the ice cover thickness at Lake Hoare is the extent of summer melting, consistent with the rising lake levels.

Wharton, Robert A., Jr.↗
NASA NTRS1992

Lunar-derived titanium alloys for hydrogen storage

Hydrogen gas, which plays an important role in many projected lunar power systems and industrial processes, can be stored in metallic titanium and in certain titanium alloys as an interstitial hydride compound. Storing and retrieving hydrogen with titanium-iron alloy requires substantially less energy investment than storage by liquefaction. Metal hydride storage systems can be designed to operate at a wide range of temperatures and pressures. A few such systems have been developed for terrestrial applications. A drawback of metal hydride storage for lunar applications is the system's large mass per mole of hydrogen stored, which rules out transporting it from earth. The transportation problem can be solved by using native lunar materials, which are rich in titanium and iron.

Love, S.↗
NASA NTRS1992

Physical properties of concrete made with Apollo 16 lunar soil sample

This paper describes the first phase of the long-term investigation for the construction of concrete lunar bases. In this phase, petrographic and scanning electron microscope examinations showed that the morphology and elemental composition of the lunar soil made it suitable for use as a fine aggregate for concrete. Based on this finding, calcium aluminate cement and distilled water were mixed with the lunar soil to fabricate test specimens. The test specimens consisted of a 1-in cube, a 1/2-in cube, and three 0.12 x 0.58 x 3.15-in beam specimens. Tests were performed on these specimens to determine compressive strength, modulus of rupture, modulus of elasticity, and thermal coefficient of expansion. Based on examination of the material and test results, it is concluded that lunar soil can be used as a fine aggregate for concrete.

Lin, T. D.↗
NASA NTRS1991

The design of a common lunar lander

The Austin Cynthesis Corporation was formed to respond to a Request for Proposal for the design of a Common Lunar Lander (CLL) capable of carrying lightweight (less than 500 kg), unspecified payload to the moon. This Final Design Report Document includes information on the requirements for the design project; the ideas proposed as solutions to the design problem; the work which has been completed in support of the design effort; justifications, validations, and verifications of decisions made during the project; and suggestions for future work to be done in support of the project. A project schedule, including current status of the items included on the schedule, as well as cost and management summaries is also included.

Driggers, Dan↗
NASA NTRS1992

Arc jet testing in NASA Ames Research Center thermophysics facilities

The Arc Jet Complex facilities at NASA Ames and their performance capabilities and support systems are presented. An overview of the typical testing procedures is provided. Attention is focused on a basic understanding of the types of facilities available at Ames for aerothermodynamic testing.

Balter-Peterson, Aliza↗
NASA NTRS1992

Arc-jet tests of coated carbon-carbon materials

The testing of coated carbon-carbon materials at high temperature in a nonequilibrium arc-driven environment is examined. Material specimens were exposed to high energy supersonic air exhausting from the NASA-Ames Research Center's 20-MW arc-jet facility. The materials were heated to required temperatures with arc-heated air for specified lengths of time. A total of 23 models were tested in two groups. In the first group of 19 samples, there were several samples that survived low test conditions and several more that survived intermediate test conditions that provided surface temperatures up to 3850 deg Rankine, but none survived when the surface temperature exceeded 4000 deg Rankine. The last four test samples with their higher surface temperatures did not fully survive either if the survival criterion is that the material must be reusable after the test.

Sheldahl, Robert E.↗
NASA NTRS1992

Evaluation of 2D ceramic matrix composites in aeroconvective environments

An evaluation is conducted of a novel ceramic-matrix composite (CMC) material system for use in the aeroconvective-heating environments encountered by the nose caps and wing leading edges of such aerospace vehicles as the Space Shuttle, during orbit-insertion and reentry from LEO. These CMCs are composed of an SiC matrix that is reinforced with Nicalon, Nextel, or carbon refractory fibers in a 2D architecture. The test program conducted for the 2D CMCs gave attention to their subsurface oxidation.

Riccitiello, Salvatore R.↗
NASA NTRS1993

A ceramic matrix composite thermal protection system for hypersonic vehicles

The next generation of hypersonic vehicles (NASP, SSTO) that require reusable thermal protection systems will experience acreage surface temperatures in excess of 1100 C. More important, they will experience a more severe physical environment than the Space Shuttle due to non-pristine launching and landing conditions. As a result, maintenance, inspection, and replacement factors must be more thoroughly incorporated into the design of the TPS. To meet these requirements, an advanced thermal protection system was conceived, designated 'TOPHAT'. This system consists of a toughened outer ceramic matrix composite (CMC) attached to a rigid reusable surface insulator (RSI) which is directly bonded to the surface. The objective of this effort was to evaluate this concept in an aeroconvective environment, to determine the effect of impacts to the CMC material, and to compare the results with existing thermal protection systems.

Riccitiello, Salvatore R.↗
NASA NTRS1991

Effects of aeroconvective environments on 2D reinforced ceramic matrix composites

The effect of aeroconvective heating environment similar to that observed a spacecraft ascent or reentry from orbit, on the performance of a commercial carbon-reinforced ceramic matrix material specimens of two configurations (orthotropic and quasi-isotropic), fabricated by the Societe Europenne Propulsion (SEP) process was investigated using the NASA Ames Research Center 20 Megawatt Panel Test facility. The performance of the commercial material was compared with the SEP prepared materials. It was found that, whereas the quasi-isotropic SEP specimens exhibited a much higher mass loss rate and a significant dimensional change upon exposure to the thermal environment than did the orthotropic ones, the commercial SEP-like materials did not exhibit these characteristics. There was no greater mass loss rate for the quasi-isotropic specimens, and no dimension changes were observed. The Nicalon reinforced materials in both configurations, as fabricated by SEP or by the commercial source, showed no mass changes and no dimensional changes.

Riccitiello, Salvatore R.↗