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At least 145 records · Page 8
All sourcesNASA NTRSDOE OSTI
NASA NTRS2016

Quasi-Periodic Fluctuations and Chromospheric Evaporation in a Solar Flare Ribbon Observed by Hinode/EIS, IRIS, and RHESSI

The Hinode/Extreme-ultraviolet Imaging Spectrometer (EIS) obtained rapid cadence (11.2 s) EUV stare spectra of an M7.3 flare ribbon in AR 12036 on 2014 April 18. Quasi-periodic (P approx. = 75.6 +/- 9.2 s) intensity fluctuations occurred in emission lines of O IV, Mg VI, Mg VII, Si VII, Fe XIV, and Fe XVI during the flare's impulsive rise, and ended when the maximum intensity in Fe XXIII was reached. The profiles of the O IV- Fe XVI lines reveal that they were all redshifted during most of the interval of quasi-periodic intensity fluctuations, while the Fe XXIII profile revealed multiple components including one or two highly blueshifted ones. This indicates that the flare underwent explosive chromospheric evaporation during its impulsive rise. Fluctuations in the relative Doppler velocities were seen, but their amplitudes were too subtle to extract significant quasi-periodicities. RHESSI detected 25-100 keV hard-X-ray sources in the ribbon near the EIS slit's pointing position during the peaks in the EIS intensity fluctuations. The observations are consistent with a series of energy injections into the chromosphere by nonthermal particle beams. Electron densities derived with Fe XIV (4.6 x 10(exp 10) per cu cm) and Mg VII (7.8 x 10(exp 9) per cu cm) average line intensity ratios during the interval of quasi-periodic intensity fluctuations, combined with the radiative loss function of an optically thin plasma, yield radiative cooling times of 32 s at 2.0 x 10(exp 6) K, and 46 s at 6.3 x 10(exp 5) K (about half the quasi-period); assuming Fe XIV's density for Fe XXIII yields a radiative cooling time of 10(exp 3) s (13 times the quasi-period) at 1.4 x 10(exp 7) K.

Brosius, Jeffrey W.↗
NASA NTRS1990

Preliminary thermal/structural analysis of a carbon-carbon/refractory-metal heat-pipe-cooled wing leading edge

This study presents preliminary thermal/structural analyses of a carbon-carbon/refractory-metal heat-pipe-cooled wing leading edge concept designed for an air breathing single-stage-to-orbit hypersonic vehicle. The concept features chordwise (i.e., normal to the leading edge) and spanwise (i.e., parallel to the leading edge) refractory-metal heat pipes which are completely embedded within a carbon-carbon primary structure. Studies of the leading edge were performed using nonlinear thermal and linear structural three-dimensional finite element analyses. The concept was shown to be thermally feasible within the limits of the assumptions made in the analyses when internal radiative cooling is present during ascent, and a three-dimensional carbon-carbon architecture is used. In addition, internal radiative cooling was found not to be necessary during descent. The linear stress analysis indicated excessively large thermal stresses in the rafractory metal walls of the heat pipes even though a soft layer of carbon was included between the heat pipe and the carbon-carbon structure in an attempt to reduce the thermal stresses. A nonlinear structural analysis may be necessary to properly model the response of the refractory-metal heat pipes.

Glass, David E.↗
NASA NTRS2016

Modeling and Simulation of Radiative Compressible Flows in Aerodynamic Heating Arc-Jet Facility

Numerical simulations of an arc heated flow inside NASA's 20 [MW] Aerodynamics heating facility (AHF) are performed in order to investigate the three-dimensional swirling flow and the current distribution inside the wind tunnel. The plasma is considered in Local Thermodynamics Equilibrium(LTE) and is composed of Air-Argon gas mixture. The governing equations are the Navier-Stokes equations that include source terms corresponding to Joule heating and radiative cooling. The former is obtained by solving an electric potential equation, while the latter is calculated using an innovative massively parallel ray-tracing algorithm. The fully coupled system is closed by the thermodynamics relations and transport properties which are obtained from Chapman-Enskog method. A novel strategy was developed in order to enable the flow solver and the radiation calculation to be preformed independently and simultaneously using a different number of processors. Drastic reduction in the computational cost was achieved using this strategy. Details on the numerical methods used for space discretization, time integration and ray-tracing algorithm will be presented. The effect of the radiative cooling on the dynamics of the flow will be investigated. The complete set of equations were implemented within the COOLFluiD Framework. Fig. 1 shows the geometry of the Anode and part of the constrictor of the Aerodynamics heating facility (AHF). Fig. 2 shows the velocity field distribution along (x-y) plane and the streamline in (z-y) plane.

Pasma flows↗
NASA NTRS2018

TPSAS-NF1676L-30053-DND

The 2016-17 Arctic sea ice growth season (October-March) exhibited one of the lowest end-of-season sea ice volume and extent of any year since 1979. An analysis of MERRA2 atmospheric reanalysis data and CERES radiative flux data reveals that a record warm and moist Arctic atmosphere supported the reduced sea ice growth through two pathways. First, numerous regional episodes of increased atmospheric temperature and moisture, transported from lower latitudes, increased the cumulative energy input from downwelling longwave surface fluxes. Second, in those same episodes, the efficiency that the atmosphere cooled radiatively to space was reduced, increasing the amount of energy retained in the Arctic atmosphere and reradiated back toward the surface. Overall, the Arctic radiative cooling efficiency shows a decreasing trend since 2000. The results presented highlight the increasing importance of atmospheric forcing on sea ice variability demonstrating that episodic Arctic atmospheric rivers, regions of elevated poleward water vapor transport, and the subsequent surface energy budget response is a critical mechanism actively contributing to the evolution of Arctic sea ice.

Bradley Hegyi↗
NASA NTRS1991

Large-scale gas dynamical processes affecting the origin and evolution of gaseous galactic halos

Observations of galactic halo gas are consistent with an interpretation in terms of the galactic fountain model in which supernova heated gas in the galactic disk escapes into the halo, radiatively cools and forms clouds which fall back to the disk. The results of a new study of several large-scale gas dynamical effects which are expected to occur in such a model for the origin and evolution of galactic halo gas will be summarized, including the following: (1) nonequilibrium absorption line and emission spectrum diagnostics for radiatively cooling halo gas in our own galaxy, as well the implications of such absorption line diagnostics for the origin of quasar absorption lines in galactic halo clouds of high redshift galaxies; (2) numerical MHD simulations and analytical analysis of large-scale explosions ad superbubbles in the galactic disk and halo; (3) numerical MHD simulations of halo cloud formation by thermal instability, with and without magnetic field; and (4) the effect of the galactic fountain on the galactic dynamo.

Shapiro, Paul R.↗
NASA NTRS2022

Impact of Clouds and Blowing Snow on Surface and Atmospheric Boundary Layer Properties over Dome C, Antarctica

Clouds and blowing snow (BLSN) occur frequently over Antarctica, where it is critical to understand their feedbacks to surface and atmospheric boundary layer processes. Dome C, an elevated East Antarctic station, dominated by lengthy periods of surface longwave (LW) radiative cooling, is selected to reveal cloud and BLSN impacts within a largely stable environment. The sky condition is classified as clear, cloudy, or BLSN, using 3 years of Cloud-Aerosol Lidar and Infrared Pathfinder Satellite Observations satellite data. Co-located and contemporaneous in situ observations are used to investigate the relationship of sky condition with surface and atmospheric boundary layer thermal structure, focusing on seasonal variability. Results show that increased downwelling LW radiation from clouds abate surface radiative cooling losses, contributing to warming during all seasons. An increase of 3°C in the mean surface air temperature is observed during spring, whereas, a more dramatic rise (around 10°C), due to accompanying large-scale subsidence, is observed during fall and winter in association with clouds. For all seasons, the wind speed and wind speed shear are strongest during BLSN events, and the surface-based inversion is weakened by cooling which peaks in a shallow above-surface turbulent layer. The stronger background stability during fall and winter seasons, restricts turbulence and BLSN depths generally to the lowest tens of meters. The Earth's cryosphere is among the most rapidly evolving yet least well-observed regions, and knowledge of clouds and BLSN interactions with the typical stable atmospheric boundary layer can help further understand energy and moisture exchanges.

CALIPSO↗
NASA NTRS1998

Contamination Control Considerations for the Next Generation Space Telescope (NGST)

The NASA Space Science Program, in its ongoing mission to study the universe, has begun planning for a telescope that will carry on the Hubble Space Telescope's exploration. This telescope, the 'Next Generation Space Telescope' (NGST), will be 6-8 meters in diameter, will be radiatively cooled to 30-60 Kelvin in order to enable extremely deep exposures at near infrared wavelengths, and will operate for a lifetime of 5-10 years. The requirement will be to measure wavelengths from 1-5 microns, with a goal to measure wavelengths from 0.6-30 microns. As such, NGST will present a new contamination control challenge. The Goddard Space Flight Center (GSFC) performed one of three preliminary feasibility studies for the NGST, presenting a telescope with an 8 meter, deployable primary mirror and a deployable secondary mirror. The telescope would be radiatively cooled, with the optical telescope assembly (OTA) and the science instrument module (SIM) isolated from the warmer spacecraft support module (SSM). The OTA and the SIM would also be shielded from sunlight with an enormous, inflatable sun-shield. The GSFC telescope was designed for launch on an Atlas HAS, which would require launching the telescope in a stowed configuration, with the SSM, antennae, sun-shield, primary mirror 'petals', and secondary mirror deployed once on-orbit. The launch configuration and deployment scenario of an exposed telescope measuring near infrared and cooled to 30-60 K are the factors presenting contamination hazards to the NGST mission. Preliminary science requirements established are: less than 20% reflectance decrease on optical surfaces over the wavelength range, and less than 0.3% obscuration of optical surfaces. In order to meet these requirements, NGST must be built and launched with careful attention to contamination control. Initial contamination control design options include strict selecting of materials and baking out of hardware down to the component level, minimizing or eliminating exposure of the OTA to sunlight or earth albedo during deployment and early on-orbit operations, cleaning of the primary and secondary mirrors at the launch site, cleaning of the launch vehicle fairing, locating thrusters and vents on the warm side of the sun shield only, and the possibility of including a deployable cover if that is shown to be necessary.

Wooldridge, Eve M.↗
NASA NTRS1989

The formation flare loops by magnetic reconnection and chromospheric ablation

Noncoplanar compressible reconnection theory is combined here with simple scaling arguments for ablation and radiative cooling to predict average properties of hot and cool flare loops as a function of the coronal vector magnetic field. For a coronal field strength of 100 G, the temperature of the hot flare loops decreases from 1.2 x 10 to the 7th K to 4.0 x 10 to the 6th K as the component of the coronal magnetic field perpendicular to the plane of the loops increases from 0 percent to 86 percent of the total field. When the perpendicular component exceeds 86 percent of the total field or when the altitude of the reconnection site exceeds 10 to the 6th km, flare loops no longer occur. Shock-enhanced radiative cooling triggers the formation of cool H-alpha flare loops with predicted densities of roughly 10 to the 13th/cu cm, and a small gap of roughly 1000 km is predicted to exist between the footpoints of the cool flare loops and the inner edges of the flare ribbons.

Forbes, T. G.↗
NASA NTRS1995

Yohkoh/SXT soft x-ray observations of sudden mass loss from the solar corona

Direct X-ray observations allow us to estimate the hot coronal mass before and after a flare or other disturbance of the type leading to a coronal mass ejection. The sudden disappearance of a large coronal structure (scale greater than 105 km) gives evidence that an ejection has occurred, if the time scales are much shorter than the conductive or radiative cooling times for such structures. A flare also typically adds large amounts of new material to the corona via evaporation resulting from the coronal energy release. This provides a competing mechanism that makes the estimation of the total mass loss somewhat difficult. We note that the X-ray observations have the advantage of covering the entire corona rather than the limb regions unlike the coronagraph observations. We have identified two examples of coronal mass disappearances. before and during long duration flare events on 21 Feb. 1992 (on the E limb) and 13 Nov. 1994 (near disk center). In latter case the total mass amounted to some 4 x 10(exp 14) g with a density of 3 x 10(exp 8)cm(exp -3) and a temperature of 2.8 MK before its disappearance. This corresponds to a radiative cooling time of some 104 S. much longer than the observed time of disappearance. We therefore suggest that these sudden mass disappearances correspond with coronal mass ejections (CMEs), and suggest that further data analysis will be able to confirm this by comparison with optical observations of specific CMEs.

Hudson, H. S.↗
NASA NTRS2014

Simulation for Prediction of Entry Article Demise (SPEAD): an Analysis Tool for Spacecraft Safety Analysis and Ascent/Reentry Risk Assessment

For the purpose of performing safety analysis and risk assessment for a probable offnominal suborbital/orbital atmospheric reentry resulting in vehicle breakup, a synthesis of trajectory propagation coupled with thermal analysis and the evaluation of node failure is required to predict the sequence of events, the timeline, and the progressive demise of spacecraft components. To provide this capability, the Simulation for Prediction of Entry Article Demise (SPEAD) analysis tool was developed. This report discusses the capabilities, modeling, and validation of the SPEAD analysis tool. SPEAD is applicable for Earth or Mars, with the option for 3 or 6 degrees-of-freedom (DOF) trajectory propagation. The atmosphere and aerodynamics data are supplied in tables, for linear interpolation of up to 4 independent variables. The gravitation model can include up to 20 zonal harmonic coefficients. The modeling of a single motor is available and can be adapted to multiple motors. For thermal analysis, the aerodynamic radiative and free-molecular/continuum convective heating, black-body radiative cooling, conductive heat transfer between adjacent nodes, and node ablation are modeled. In a 6- DOF simulation, the local convective heating on a node is a function of Mach, angle-ofattack, and sideslip angle, and is dependent on 1) the location of the node in the spacecraft and its orientation to the flow modeled by an exposure factor, and 2) the geometries of the spacecraft and the node modeled by a heating factor and convective area. Node failure is evaluated using criteria based on melting temperature, reference heat load, g-load, or a combination of the above. The failure of a liquid propellant tank is evaluated based on burnout flux from nucleate boiling or excess internal pressure. Following a component failure, updates are made as needed to the spacecraft mass and aerodynamic properties, nodal exposure and heating factors, and nodal convective and conductive areas. This allows the trajectory to be propagated seamlessly in a single run, inclusive of the trajectories of components that have separated from the spacecraft. The node ablation simulates the decreasing mass and convective/reference areas, and variable heating factor. A built-in database provides the thermo-mechanical properties of For the purpose of performing safety analysis and risk assessment for a probable offnominal suborbital/orbital atmospheric reentry resulting in vehicle breakup, a synthesis of trajectory propagation coupled with thermal analysis and the evaluation of node failure is required to predict the sequence of events, the timeline, and the progressive demise of spacecraft components. To provide this capability, the Simulation for Prediction of Entry Article Demise (SPEAD) analysis tool was developed. This report discusses the capabilities, modeling, and validation of the SPEAD analysis tool. SPEAD is applicable for Earth or Mars, with the option for 3 or 6 degrees-of-freedom (DOF) trajectory propagation. The atmosphere and aerodynamics data are supplied in tables, for linear interpolation of up to 4 independent variables. The gravitation model can include up to 20 zonal harmonic coefficients. The modeling of a single motor is available and can be adapted to multiple motors. For thermal analysis, the aerodynamic radiative and free-molecular/continuum convective heating, black-body radiative cooling, conductive heat transfer between adjacent nodes, and node ablation are modeled. In a 6- DOF simulation, the local convective heating on a node is a function of Mach, angle-ofattack, and sideslip angle, and is dependent on 1) the location of the node in the spacecraft and its orientation to the flow modeled by an exposure factor, and 2) the geometries of the spacecraft and the node modeled by a heating factor and convective area. Node failure is evaluated using criteria based on melting temperature, reference heat load, g-load, or a combination of the above. The failure of a liquid propellant tank is evaluated based on burnout flux from nucleate boiling or excess internal pressure. Following a component failure, updates are made as needed to the spacecraft mass and aerodynamic properties, nodal exposure and heating factors, and nodal convective and conductive areas. This allows the trajectory to be propagated seamlessly in a single run, inclusive of the trajectories of components that have separated from the spacecraft. The node ablation simulates the decreasing mass and convective/reference areas, and variable heating factor. A built-in database provides the thermo-mechanical properties of

Ling, Lisa↗
NASA NTRS1976

Liquid rocket engine nozzles

The nozzle is a major component of a rocket engine, having a significant influence on the overall engine performance and representing a large fraction of the engine structure. The design of the nozzle consists of solving simultaneously two different problems: the definition of the shape of the wall that forms the expansion surface, and the delineation of the nozzle structure and hydraulic system. This monography addresses both of these problems. The shape of the wall is considered from immediately upstream of the throat to the nozzle exit for both bell and annular (or plug) nozzles. Important aspects of the methods used to generate nozzle wall shapes are covered for maximum-performance shapes and for nozzle contours based on criteria other than performance. The discussion of structure and hydraulics covers problem areas of regeneratively cooled tube-wall nozzles and extensions; it treats also nozzle extensions cooled by turbine exhaust gas, ablation-cooled extensions, and radiation-cooled extensions. The techniques that best enable the designer to develop the nozzle structure with as little difficulty as possible and at the lowest cost consistent with minimum weight and specified performance are described.

Source record↗
NASA NTRS1977

Liquid rocket engine self-cooled combustion chambers

Self-cooled combustion chambers are chambers in which the chamber wall temperature is controlled by methods other than fluid flow within the chamber wall supplied from an external source. In such chambers, adiabatic wall temperature may be controlled by use of upstream fluid components such as the injector or a film-coolant ring, or by internal flow of self-contained materials; e.g. pyrolysis gas flow in charring ablators, and the flow of infiltrated liquid metals in porous matrices. Five types of self-cooled chambers are considered in this monograph. The name identifying the chamber is indicative of the method (mechanism) by which the chamber is cooled, as follows: ablative; radiation cooled; internally regenerative (Interegen); heat sink; adiabatic wall. Except for the Interegen and heat sink concepts, each chamber type is discussed separately. A separate and final section of the monograph deals with heat transfer to the chamber wall and treats Stanton number evaluation, film cooling, and film-coolant injection techniques, since these subjects are common to all chamber types. Techniques for analysis of gas film cooling and liquid film cooling are presented.

Source record↗
NASA NTRS1978

Design and fabrication of a radiative actively cooled honeycomb sandwich structural panel for a hypersonic aircraft

The panel assembly consisted of an external thermal protection system (metallic heat shields and insulation blankets) and an aluminum honeycomb structure. The structure was cooled to temperature 442K (300 F) by circulating a 60/40 mass solution of ethylene glycol and water through dee shaped coolant tubes nested in the honeycomb and adhesively bonded to the outer skin. Rene'41 heat shields were designed to sustain 5000 cycles of a uniform pressure of + or - 6.89kPa (+ or - 1.0 psi) and aerodynamic heating conditions equivalent to 136 kW sq m (12 Btu sq ft sec) to a 422K (300 F) surface temperature. High temperature flexible insulation blankets were encased in stainless steel foil to protect them from moisture and other potential contaminates. The aluminum actively cooled honeycomb sandwich structural panel was designed to sustain 5000 cycles of cyclic in-plane loading of + or - 210 kN/m (+ or - 1200 lbf/in.) combined with a uniform panel pressure of + or - 6.89 kPa (?1.0 psi).

Ellis, D. A.↗
NASA NTRS1985

Convoluted nozzle design for the RL10 derivative 2B engine

The convoluted nozzle is a conventional refractory metal nozzle extension that is formed with a portion of the nozzle convoluted to show the extendible nozzle within the length of the rocket engine. The convoluted nozzle (CN) was deployed by a system of four gas driven actuators. For spacecraft applications the optimum CN may be self-deployed by internal pressure retained, during deployment, by a jettisonable exit closure. The convoluted nozzle is included in a study of extendible nozzles for the RL10 Engine Derivative 2B for use in an early orbit transfer vehicle (OTV). Four extendible nozzle configurations for the RL10-2B engine were evaluated. Three configurations of the two position nozzle were studied including a hydrogen dump cooled metal nozzle and radiation cooled nozzles of refractory metal and carbon/carbon composite construction respectively.

Source record↗