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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 19 records
All sourcesNASA NTRSDOE OSTI
NASA NTRS1998

Diffusive Gas Loss from Silica Glass Ampoules at Elevated Temperatures

Changes in the pressure of hydrogen, helium and neon due to diffusion through the wall of silica crystal growth ampoules at elevated temperatures were determined experimentally. We show that, while both He- and Ne-losses closely follow conventional model of diffusive gas permeation through the wall, hydrogen losses, in particular at low fill pressures, can be much larger. This is interpreted in terms of the high solubility of hydrogen in silica glasses.

Palosz, W.↗
NASA NTRS1985

Measurement of energy distribution in flowing hydrogen microwave plasmas

An electrothermal propulsion concept utilizing a microwave plasma system as the mechanism to convert electromagnetic energy into kinetic energy of a flowing gas is investigated. A calorimetry system enclosing a microwave plasma system has been developed to accurately measure the energy inputs and outputs of the microwave plasma system. The rate of energy transferred to the gas can be determined to within + or - 1.8 W from an energy balance around the microwave plasma system. The percentage of the power absorbed by the microwave plasma system transferred to the hydrogen gas as it flows through the system is found to increase with the increasing flow rate, to decrease with the increasing pressure, and to be independent of the absorbed power. An upper bound for the hydrogen gas temperature is estimated from the energy content, heat capacity, and flow rate of the gas stream. A lower bound for an overall heat-transfer coefficient is then calculated, characterizing the energy loss from the hydrogen gas stream to the air cooling of the plasma discharge tube wall. The heat-transfer coefficient is found to increase with the increasing flow rate and pressure and to be independent of the absorbed power. This result indicates that a convective-type mechanism is responsible for the energy transfer.

Chapman, R.↗
NASA NTRS2013

Prediction of Ablation Rates from Solid Surfaces Exposed to High Temperature Gas Flow

A mathematical model and a solution algorithm is developed to study the physics of high temperature heat transfer and material ablation and identify the problems associated with the flow of hydrogen gas at very high temperatures and velocities through pipes and various components of Nuclear Thermal Rocket (NTR) motors. Ablation and melting can be experienced when the inner solid surface of the cooling channels and the diverging-converging nozzle of a Nuclear Thermal Rocket (NTR) motor is exposed to hydrogen gas flow at temperatures around 2500 degrees Kelvin and pressures around 3.4 MPa. In the experiments conducted on typical NTR motors developed in 1960s, degradation of the cooling channel material (cracking in the nuclear fuel element cladding) and in some instances melting of the core was observed. This paper presents the results of a preliminary study based on two types of physics based mathematical models that were developed to simulate the thermal-hydrodynamic conditions that lead to ablation of the solid surface of a stainless steel pipe exposed to high temperature hydrogen gas near sonic velocities. One of the proposed models is one-dimensional and assumes the gas flow to be unsteady, compressible and viscous. An in-house computer code was developed to solve the conservations equations of this model using a second-order accurate finite-difference technique. The second model assumes the flow to be three-dimensional, unsteady, compressible and viscous. A commercial CFD code (Fluent) was used to solve the later model equations. Both models assume the thermodynamic and transport properties of the hydrogen gas to be temperature dependent. In the solution algorithm developed for this study, the unsteady temperature of the pipe is determined from the heat equation for the solid. The solid-gas interface temperature is determined from an energy balance at the interface which includes heat transfer from or to the interface by conduction, convection, radiation, and ablation. Two different ablation models are proposed to determine the heat loss from the solid surface due to the ablation of the solid material. Both of them are physics based. Various numerical simulations were carried out using both models to predict the temperature distribution in the solid and in the gas flow, and then predict the ablation rates at a typical NTR motor hydrogen gas temperature and pressure. Solid mass loss rate per foot of a pipe was also calculated from these predictions. The results are presented for fully developed turbulent flow conditions in a sample SS pipe with a 6 inch diameter.

Akyuzlu, Kazim M.↗
NASA NTRS1979

Initial comparison of single cylinder Stirling engine computer model predictions with test results

A NASA developed digital computer code for a Stirling engine, modelling the performance of a single cylinder rhombic drive ground performance unit (GPU), is presented and its predictions are compared to test results. The GPU engine incorporates eight regenerator/cooler units and the engine working space is modelled by thirteen control volumes. The model calculates indicated power and efficiency for a given engine speed, mean pressure, heater and expansion space metal temperatures and cooler water inlet temperature and flow rate. Comparison of predicted and observed powers implies that the reference pressure drop calculations underestimate actual pressure drop, possibly due to oil contamination in the regenerator/cooler units, methane contamination in the working gas or the underestimation of mechanical loss. For a working gas of hydrogen, the predicted values of brake power are from 0 to 6% higher than experimental values, and brake efficiency is 6 to 16% higher, while for helium the predicted brake power and efficiency are 2 to 15% higher than the experimental.

Tew, R. C., Jr.↗
NASA NTRSDate not supplied

Hot hydrogen testing of carbides developed for nuclear thermal propulsion

Carbides are candidate materials for use in nuclear thermal propulsion, either as the primary material in nuclear fuel elements or as protective coatings. During operation, the fuel elements are exposed to flowing hydrogen gas at high temperature and it is important that fuel mass loss resulting from hydrogen corrosion be kept to a minimum. In this study, ZrC and NbC were fully consolidated by spark plasma sintering at 2100 °C, and relative densities greater than 98% were achieved. The samples were characterized by scanning electron microscopy and x-ray diffraction before and after hydrogen testing. The samples were exposed to hot flowing hydrogen gas at a pressure of 1 atm., at a temperature of 2500 K, and for durations up to 3 hours. It was found that the mass loss rate decreased with increasing exposure, before reaching a steady state value of about 1 mg/m2s. Oxygen in the samples increased the mass loss rate. Samples for which oxygen could be detected by x-ray diffraction had their mass loss rate increase by a factor of 10. Overall, the mass loss rates compare quite favorably to those previously reported.

M. P. Volz↗
NASA NTRS1972

A 10,000-gpm liquid hydrogen transfer system for the Saturn/Apollo program.

Brief description of the design and operation of the liquid hydrogen transfer system used to service the Saturn V launch vehicle. The cryogenic loading of the huge booster begins eight hours before the scheduled liftoff. The first three hours of fueling are spent in cold hydrogen gas conditioning of the fuel tank. The cold hydrogen gas is provided by vaporizing liquid hydrogen from the storage tank and routing the resultant gas through the fill system. Boil-off losses after loading are continuously replaced through control valves which are driven by a computer system. The liquid hydrogen transfer system is made up of a number of subsystems including the 850,000 gal storage tank whose boil-off losses amount to only 200 gal/day, the pressurization system, the burn pond for controlled disposal of hydrogen waste gas, the storage tank fill manifold, and the hazardous gas monitoring system. Some of the subsystems and components are redundant to provide a high degree of reliability.

Wybranowski, E., Jr.↗
NASA NTRS1976

Laser-heated rocket studies

CW laser heated rocket propulsion was investigated in both the flowing core and stationary core configurations. The laser radiation considered was 10.6 micrometers, and the working gas was unseeded hydrogen. The areas investigated included initiation of a hydrogen plasma capable of absorbing laser radiation, the radiation emission properties of hot, ionized hydrogen, the flow of hot hydrogen while absorbing and radiating, the heat losses from the gas and the rocket performance. The stationary core configuration was investigated qualitatively and semi-quantitatively. It was found that the flowing core rockets can have specific impulses between 1,500 and 3,300 sec. They are small devices, whose heating zone is only a millimeter to a few centimeters long, and millimeters to centimeters in radius, for laser power levels varying from 10 to 5,000 kW, and pressure levels of 3 to 10 atm. Heat protection of the walls is a vital necessity, though the fraction of laser power lost to the walls can be as low as 10% for larger powers, making the rockets thermally efficient.

Kemp, N. H.↗
NASA NTRS2004

Polymer/Silicate Nanocomposites Used to Manufacture Gas Storage Tanks With Reduced Permeability

Over the past decade, there has been considerable research in the area of polymer-layered silicate nanocomposites. This research has shown that the dispersion of small amounts of an organically modified layered silicate improves the polymer strength, modulus, thermal stability, and barrier properties. There have been several reports on the dispersion of layered silicates in an epoxy matrix. Potential enhancements to the barrier properties of epoxy/silicate nanocomposites make this material attractive for low permeability tankage. Polymer matrix composites (PMCs) have several advantages for cryogenic storage tanks. They are lightweight, strong, and stiff; therefore, a smaller fraction of a vehicle's potential payload capacity is used for propellant storage. Unfortunately, the resins typically used to make PMC tanks have higher gas permeability than metals. This can lead to hydrogen loss through the body of the tank instead of just at welds and fittings. One approach to eliminate this problem is to build composite tanks with thin metal liners. However, although these tanks provide good permeability performance, they suffer from a substantial mismatch in the coefficient of thermal expansion, which can lead to failure of the bond between the liner and the body of the tank. Both problems could be addressed with polymersilicate nanocomposites, which exhibit reduced hydrogen permeability, making them potential candidates for linerless PMC tanks. Through collaboration with Northrop Grumman and Michigan State University, nanocomposite test tanks were manufactured for the NASA Glenn Research Center, and the helium permeability was measured. An organically modified silicate was prepared at Michigan State University and dispersed in an epoxy matrix (EPON 826/JeffamineD230). The epoxy/silicate nanocomposites contained either 0 or 5 wt% of the organically modified silicate. The tanks were made by filament winding carbon fibers with the nanocomposite resin. Helium permeability was measured by Northrop Grumman, showing that the leak rate/day of the nanocomposite matrix tank was approximately 80-percent less than that of the neat epoxy matrix tank.

Campbell, Sandi G.↗
NASA NTRS1992

Capacity retention in hydrogen storage alloys

Results of our examination of the properties of several candidate materials for hydrogen storage electrodes and their relation to the decrease in H-storage capacity upon open-circuit storage over time are reported. In some of the alloy samples examined to date, only about 10 percent of the hydrogen capacity was lost upon storage for 20 days, while in others, this number was as high as 30 percent for the same period of time. This loss in capacity is attributed to two separate mechanisms: (1) hydrogen desorbed from the electrode due to pressure differences between the cell and the electrode sample; and (2) chemical and/or electrochemical degradation of the alloy electrode upon exposure to the cell environment. The former process is a direct consequence of the equilibrium dissociation pressure of the hydride alloy phase and the partial pressure of hydrogen in the hydride phase in equilibrium with that in the electrolyte environment, while the latter is related to the stability of the alloy phase in the cell environment. Comparison of the equilibrium gas-phase dissociation pressures of these alloys indicate that reversible loss of hydrogen capacity is higher in alloys with P(eqm) greater than 1 atm than in those with P(eqm) less than 1 atm.

Anani, A.↗
NASA NTRS1982

Advanced turbine study

Experiments to determine the available increase in turbine horsepower achieved by increasing turbine inlet temperature over a range of 1800 to 2600 R, while applying current gas turbine airfoil cling technology are discussed. Four cases of rocket turbine operating conditions were investigated. Two of the cases used O2/H2 propellant, one with a fuel flowrate of 160 pps, the other 80 pps. Two cases used O2/CH4 propellant, each having different fuel flowrates, pressure ratios, and inlet pressures. Film cooling was found to be the required scheme for these rocket turbine applications because of the high heat flux environments. Conventional convective or impingement cooling, used in jet engines, is inadequate in a rocket turbine environment because of the resulting high temperature gradients in the airfoil wall, causing high strains and low cyclic life. The hydrogen-rich turbine environment experienced a loss, or no gain, in delivered horsepower as turbine inlet temperature was increased at constant airfoil life. The effects of film cooling with regard to reduced flow available for turbine work, dilution of mainstream gas temperature and cooling reentry losses, offset the relatively low specific work capability of hydrogen when increasing turbine inlet temperature over the 1800 to 2600 R range. However, the methane-rich environment experienced an increase in delivered horsepower as turbine inlet temperature was increased at constant airfoil life. The results of a materials survey and heat transfer and durability analysis are discussed.

Source record↗
NASA NTRS1990

Characterization of hydrogen interactions with the surface of alpha-2 titanium aluminide alloys

The interaction of hydrogen at the surface of alpha-2 titanium aluminide alloys with Nb additions has been investigated with Auger electron spectroscopy (AES) and reflected electron energy loss (REELS) spectroscopy. The hydrogen was either adsorbed onto the alloy surface from the gas phase or it was pre-charged into the alloy forming an alpha-2 hydride. No effect of either adsorbed or absorbed hydrogen was observed in the Auger spectra. A shift in the bulk plasmon loss energy (observed with REELS) was measured for both conditions. The potential significance of these results in relation to hydrogen interactions in this alloy is discussed.

Shanabarger, Mickey R.↗
NASA NTRS1968

Nuclear-Rocket Propulsion

The three basic types of nuclear power-plants (solid, liquid, and gas core) are compared on the bases of performance potential and the status of current technology. The solid-core systems are expected to have impulses in the range of 850 seconds, any thrust level (as long as it is greater than 10,000 pounds (44,480 newtons)), and thrust-to-engine-weight ratios of 2 to 20 pounds per pound (19.7 to 197 newtons per kilogram). There is negligible or no fuel loss from the solid-core system. The solid-core system, of course, has had the most work done on it. Large-scale tests have been performed on a breadboard engine that has produced specific impulses greater than 700 seconds at thrust levels of about 50,000 pounds (222,000 newtons). The liquid-core reactor would be interesting in the specific impulse range of 1200 to 1500 seconds. Again, any thrust level can be obtained depending on how big or small the reactor is made. The thrust-to-engine weight ratio for these systems would be in the range of 1 to 10. The discouraging feature of the liquid-core system is the high fuel-loss ratio anticipated. Values of 0.01 to 0.1 pound (0.00454 to 0.0454 kilograms) or uranium loss per pound (0.454 kilograms) of hydrogen are expected, if impulses in the range of 1200 to 1500 seconds are desired. The gas-core reactor shows specific impulses in the range of 1500 to 2500 seconds. The thrust levels should be at least as high as the weight so that the thrust-to-weight ratio does not go below 1. Because the engine weight is not expected to be under 100,000 pounds (444,800 newtons), thrust levels higher than 100,000 pounds (448,000 newtons) are of interest. The thrust-to-engine weights, in that case, would run from 1 to 20 pounds per pound (9.8 to 19.7 kilograms). Gas-core reactors tend to be very large, and can have high thrust-to-weight ratios. As in the case of the liquid-core system, the fuel loss that will be attendant with gas cores as envisioned today will be rather high. The loss rates will be 0.01 to 0.1 pound of uranium (0.00454 to 0.0454 kilograms) for each pound (0.454 kilograms) of hydrogen.

Rom, Frank E.↗
NASA NTRS2020

A remnant planetary core in the hot-Neptune desert

The interiors of giant planets remain poorly understood. Even for the planets in the Solar System, difficulties in observation lead to large uncertainties in the properties of planetary cores. Exoplanets that have undergone rare evolutionary processes provide a route to understanding planetary interiors. Planets found in and near the typically barren hot-Neptune ‘desert’ (a region in mass–radius space that contains few planets) have proved to be particularly valuable in this regard. These planets include HD149026b, which is thought to have an unusually massive core, and recent discoveries such as LTT9779b and NGTS-4b, on which photoevaporation has removed a substantial part of their outer atmospheres. Here we report observations of the planet TOI-849b, which has a radius smaller than Neptune’s but an anomalously large mass of 39.1(+2.7−2.6) Earth masses and a density of 5.2(+0.7−0.8) grams per cubic centimetre, similar to Earth’s. Interior-structure models suggest that any gaseous envelope of pure hydrogen and helium consists of no more than 3.9(+0.8−0.9) per cent of the total planetary mass. The planet could have been a gas giant before undergoing extreme mass loss via thermal self-disruption or giant planet collisions, or it could have avoided substantial gas accretion, perhaps through gap opening or late formation. Although photoevaporation rates cannot account for the mass loss required to reduce a Jupiter-like gas giant, they can remove a small (a few Earth masses) hydrogen and helium envelope on timescales of several billion years, implying that any remaining atmosphere on TOI-849b is likely to be enriched by water or other volatiles from the planetary interior. We conclude that TOI-849b is the remnant core of a giant planet.

David J. Armstrong↗
NASA NTRS1980

Pulsation and mass loss in Mira variables

The behavior of pulsation in the outer layers of a typical Mira variable was studied in the adiabatic and isothermal limits. A shock wave propagates outward once per period and the radial velocity obtained from observations of hydrogen emission lines is identified with the velocity of gas in the post shock region. In the adiabatic case, mass loss in the form of a steady stellar wind was produced. In the isothermal case, no continuous mass loss was produced but occasional ejection of shells occur. Pulsation introduced into a star undergoing steady mass loss as a result of radiation pressure acting on grains caused the mass loss rate to increase by a factor of approximately 40, while the terminal velocity of the flow was almost unaltered.

Wood, P. R.↗
NASA NTRS1990

Integrated lunar materials manufacturing process

A manufacturing plant and process for production of oxygen on the moon uses lunar minerals as feed and a minimum of earth-imported, process materials. Lunar feed stocks are hydrogen-reducible minerals, ilmenite and lunar agglutinates occurring in numerous, explored locations mixed with other minerals in the pulverized surface layer of lunar soil known as regolith. Ilmenite (FeTiO.sub.3) and agglutinates contain ferrous (Fe.sup.+2) iron reducible by hydrogen to yield H.sub.2 O and metallic Fe at about 700.degree.-1,200.degree. C. The H.sub.2 O is electrolyzed in gas phase to yield H.sub.2 for recycle and O.sub.2 for storage and use. Hydrogen losses to lunar vacuum are minimized, with no net hydrogen (or any other earth-derived reagent) consumption except for small leaks. Feed minerals are surface-mined by front shovels and transported in trucks to the processing area. The machines are manned or robotic. Ilmenite and agglutinates occur mixed with silicate minerals which are not hydrogen-reducible at 700.degree.-1,200.degree. C. and consequently are separated and concentrated before feeding to the oxygen generation process. Solids rejected from the separation step and reduced solids from the oxygen process are returned to the mine area. The plant is powered by nuclear or solar power generators. Vapor-phase water electrolysis, a staged, countercurrent, fluidized bed reduction reactor and a radio-frequency-driven ceramic gas heater are used to improve thermal efficiency.

Gibson, Michael A.↗
NASA NTRS2023

Assessment of Gaseous Helium Cooling in a Superconducting Coil Test Rig

An experimental rig was designed and constructed at the NASA Glenn Research Center to measure the alternating current (AC) losses of round, multifilament, superconducting (SC), magnesium diboride (MgB2) wire in the temperature range from 20 to 30 K. The rig resembles a permanent magnet (PM) motor with the stator windings omitted. Experiment packs containing the superconductor samples will be placed in the annular space between the rotor and the stator back iron. The original plan was to measure the AC losses by calorimetry using nonboiling liquid hydrogen as the cooling fluid. The present plan is to use helium gas (GHe) under a few bar pressure instead. This report summarizes the calculations that indicate that reasonable results can be obtained, even though the mass density of the GHe will be only a fraction of that of liquid hydrogen. It is necessary to show that the heat transfer coefficient from SC wire to GHe can be high enough, that the temperature rise of the He between the experiment inlet and outlet can be small enough, and that sufficient He flow can be driven by a commercially available He fan. Both straight wire samples to validate loss models and coils to simulate motor stator coils can be tested in the rig. The lowest superconductor temperature that is feasible to reach in the rig is estimated. The temperature range available for samples will be wider with the gaseous He system than it would have been with liquid hydrogen.

supeconductors↗
NASA NTRS1991

National Aerospace Plane Engine Seals: High Temperature Seal Performance Evaluation

The key to the successful development of the single stage to orbit National Aerospace Plane (NASP) is the successful development of combined cycle ramjet/scramjet engines that can propel the vehicle to 17,000 mph to reach low Earth orbit. To achieve engine performance over this speed range, movable engine panels are used to tailor engine flow that require low leakage, high temperature seals around their perimeter. NASA-Lewis is developing a family of new high temperature seals to form effective barriers against leakage of extremely hot (greater than 2000 F), high pressure (up to 100 psi) flow path gases containing hydrogen and oxygen. Preventing backside leakage of these explosive gas mixtures is paramount in preventing the potential loss of the engine or the entire vehicle. Seal technology development accomplishments are described in the three main areas of concept development, test, and evaluation and analytical development.

Steinetz, Bruce M.↗