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A Monte Carlo calculation of the quark-liberating phase transition in lattice quantum chromodynamics is presented. The transition temperature as a function of the lattice coupling g does not scale according to the perturbative beta function for 6/g-squared less than 6.1. Finite-size scaling is used in analyzing the properties of the lattice system near the transition point.
The mass of a radiative actively cooled panel was compared to the mass of a bare actively cooled panel designed to the same conditions and constraints. The approach was to design and optimize a 0.61 x 6.1 m full scale panel which combines radiative and active cooling to control structural temperatures to levels compatible with use of lightweight materials and to fabricate a 0.61 x 1.22 m panel for performance testing. Results of the design and optimization of the full scale radiative actively cooled structural panel, including radiative concept selection, final configuration details, test panel description, and conclusions are summarized.
Soon after the Space Shuttle Columbia accident occurred last year, a group of CFD analysts from NASA centers and private industry was organized to help determine the cause of the accident. This group was under the direction of the Applied Aeroscience and CFD Branch of the Aeroscience and Flight Mechanics Division at the Johnson Space Center. For external flow simulations, noncommercial CFD codes that specialize in hypersonic or high Mach number flows were used. These tools were used to determine heating rates, pressures, and temperatures for a large number of vehicle damage scenarios. Lockheed Martin Space Operations was called upon to provide CFD support in the area of internal flows within the shuttle wing cavity, and or these simulations, FLUENT 6.1 was chosen. Two large-scale, simplified models were run to understand the flow patterns once a breach of the internal wing cavity was initiated. The results were primarily used to visualize flow patterns within the wing cavity. The first CFD model included the entire left wing without the wheel. well cavity. The purpose of the first model, which did not include the reinforced carbon-carbon (RCC) cavity along the wing leading edge, was to visualize the flow field within the wing cavity immediately after the leading edge spar breach, This model assumed that the flow coming into the wing cavity was nom1 to the spar. It included all o f the primary vents that allow for flow between the main cavities of the wing. A six-inch diameter hole was modeled in the spar at the approximate location where the spar breach was judged to have occurred, which was between RCC panels 8 and 9. The results of the modeling showed that at this location, the high temperature, high velocity gas stream entering the wing cavity impinged on the outboard wheel well cavity. Instrumentation in the Shuttle wheel well cavity registered abnormal temperatures during reentry, so the FLUENT results helped support the conclusion of the accident investigation team that the spar breach was in the RCC panel 8-9 area, and that the initial spar breach was likely entering the wing cavity normal to the spar. This model also showed that the flow entering the wing cavity tended to swirl within the cavity just outboard of the wheel well, and did not initially penetrate further into the rear cavities of the wing. The second CFD model was a 2-D simulation of the left wing cavity and the RCC cavity. It was used to visualize the flow through the RCC breach, through the wing spar breach, and into the wing cavity directly outboard of the wheel well. The purpose of this model was to verify whether or not it was possible for the flow to come into the wing cavity normal to the leading edge spar. The results from this 2-D model showed that the internal structure behind the RCC panels probably deflected the flow entering the RCC cavity so that it impinged normal to the spar. As in the 3-D model, this deflected flow stream was found to impinge on the wheel well outer wall. This model again supported the conclusions regarding the location of the spar breach and how the flow behaved inside the wing cavity.
Soon after the Space Shuttle Columbia accident occurred last year, a group of CFD analysts from NASA centers and private industry was organized to help determine the cause of the accident. This group was under the direction of the Applied Aeroscience and CFD Branch of the Aeroscience and Flight Mechanics Division at the Johnson Space Center. For external flow simulations, noncommercia2 CFD codes that specialize in hypersonic or high Mach number flows were used. These tools were used to determine heating rates, pressures, and temperatures for a large number of vehicle damage scenarios. Lockheed Martin Space Operations was called upon to provide CFD support in the area of internal flows within the shuttle wing cavity, and for these simulations, FLUENT 6.1 was chosen. Two large-scale, simplified models were m to understand the flow patterns once a breach of the internal wing cavity was initiated. The results were primarily used to visualize flow patterns within the wing cavity. The first CFD model included the entire lee wing without the wheel well cavity. The purpose of the first model, which did not include the reinforced carbon-carbon (RCC) cavity along the wing leading edge, was to visualize the flow field within the wing cavity immediately after the leading edge spar breach, This model assumed that the flow coming into the wing cavity was normal to the spar. It included all of the primary vents that allow for flow between the main cavities of the wing. A six-inch diameter hole was modeled in the spar at the approximate location where the spar breach was judged to have occurred, which was between RCC panels 8 and 9. The results of the modeling showed that at this location, the high temperature, high velocity gas stream entering the wing cavity impinged on the outboard wheel well cavity. Instrumentation in the Shuttle wheel well cavity registered abnormal temperatures during reentry, so the FLUENT results helped support the conclusion of the accident investigation team that the spar breach was in the RCC panel 8-9 area, and that the initial spar breach was likely entering the wing cavity normal to the spar. This model also showed that the flow entering the wing cavity tended to swirl within the cavity just outboard of the wheel well, and did not initially penetrate further into the rear cavities of the wing. The second CFD model was a 2-D simulation of the left wing cavity and the RCC cavity. It was used to visualize the flow through the RCC breach, through the wing spar breach, and into the wing cavity directly outboard of the wheel well. The purpose of this model was to verify whether or not it was possible for the flow to come into the wing cavity normal to the leading edge spar. The results from this 2-D model showed that the internal structure behind the RCC panels probably deflected the flow entering the RCC cavity so that it impinged normal to the spar. As in the 3-D model, this deflected flow stream was Found to impinge on the wheel well outer wall. This model again supported the conclusions regarding the location of the spar breach and how the flow behaved inside the wing cavity.
Measurements and predictions of the X-33 turbulent aeroheating environment have been performed at Mach 6, perfect-gas air conditions. The purpose of this investigation was to compare measured turbulent aeroheating levels on smooth models, models with discrete trips, and models with arrays of bowed panels (which simulate bowed thermal protections system tiles) with each other and with predictions from two Navier-Stokes codes, LAURA and GASP. The wind tunnel testing was conducted at free stream Reynolds numbers based on length of 1.8 x 10(exp 6) to 6.1 x 10(exp 6) on 0.0132 scale X-33 models at a = 40-deg. Turbulent flow was produced by the discrete trips and by the bowed panels at ill but the lowest Reynolds number, but turbulent flow on the smooth model was produced only at the highest Reynolds number. Turbulent aeroheating levels on each of the three model types were measured using global phosphor thermography and were found to agree to within .he estimated uncertainty (plus or minus 15%) of the experiment. Computations were performed at the wind tunnel free stream conditions using both codes. Turbulent aeroheating levels predicted using the LAURA code were generally 5%-10% lower than those from GASP, although both sets of predictions fell within the experimental accuracy of the wind tunnel data.
We present the first in-depth statistical survey of flare source heights observed by RHESSI. Flares were found using a flare-finding algorithm designed to search the 6-10 keV count-rate when RHESSI's full sensitivity was available in order to find the smallest events (Christe et al., 2008). Between March 2002 and March 2007, a total of 25,006 events were found. Source locations were determined in the 4-10 keV, 10-15 keV, and 15-30 keV energy ranges for each event. In order to extract the height distribution from the observed projected source positions, a forward-fit model was developed with an assumed source height distribution where height is measured from the photosphere. We find that the best flare height distribution is given by g (h) oc exp(-h/lambda) where lambda = 6.1 plus or minus 0.3 Mm is the scale height. A power law height distribution with a negative power law index, gamma = 3.1 plus or minus 0.3 is also consistent with the data. Interpreted as thermal loop top sources, these heights are compared to loops generated by a potential field model (PFSS). The measured flare heights distribution are found to be much steeper than the potential field loop height distribution which may be a signature of the flare energization process.
Study of the electron-impact excitation spectrum of O2 at 20 and 45 eV impact energies and at scattering angles ranging from 10 to 90 deg. The angular behavior of the differential scattering cross sections for excitation of the a super 1 delta sub g, b super 1 sigma sub g (+), B super 3 sigma sub u (-) states for the 9.97-eV ('longest' band) and the 10.29-eV ('second' band) transitions, for the broad feature at 6.1 eV energy-loss, and for elastic scattering is determined. The experimentally measured relative differential and integral cross sections for these processes are approximately normalized to the absolute scale. The intensities of the different transitions in optical and electron-impact spectra are compared, and the importance of spin-orbit coupling and exchange processes is discussed. It is found that the energy-loss feature at 6.1 eV in the electron-impact spectrum is mostly due to the excitation of the c super 1 sigma sub u (-) state, and not the A super 3 sigma sub u (+) state, as had been previously thought.
Power spectra of measurements of the magnetic field strength in the heliosphere obtained by Voyager 1 between 1 AU and 9 AU have the form of a power law f exp -a from periods of several hours to at least 6 days. The exponent was a = 2.0 + or - 0.05 for all of the spectra considered, which is the exponent for a series of steps and for Burgers' (1971) turbulence. Spectra of large-scale speed fluctuations also have the form f exp - b from a period of a few hours to periods greater than 13 days in the region from 1 AU to 8.9 AU. The exponent b is generally somewhat larger than b = 2, implying some 'persistence' of the speed fluctuations. The low-frequency cutoff (outer cutoff) of the power law increases from a period of 6.5 days at 1 AU to 26 days at (6.1-8.9) AU, which can be attributed to: (1) the coalescence of interaction regions and (2) a transfer of energy from the spectrum of large-scale speed fluctuations. The outer cutoff of the spectrum of speed fluctuations increases from a period of 13 days at 1 AU to 26 days between a few AU and 8.9 AU. Both the magnetic field strength fluctuations and the speed fluctuations have fractal behavior, suggesting that they are self affine rather than dominated by a few large discontinuities.
An experimental investigation was performed to determine the effect of pressure gradient on the heat transfer to space shuttle reusable surface insulation (RSI) tile array gaps under thick, turbulent boundary layer conditions. Heat transfer and pressure measurements were obtained on a curved array of full-scale simulated RSI tiles in a tunnel wall boundary layer at a nominal freestream Mach number of 10.3 and freestream unit Reynolds numbers of 1.6, 3.3, and and 6.1 million per meter. Transverse pressure gradients were induced over the model surface by rotating the curved array with respect to the flow. Definition of the tunnel wall boundary layer flow was obtained by measurement of boundary layer pitot pressure profiles, and flat plate wall pressure and heat transfer. Flat plate wall heat transfer data were correlated and a method was derived for prediction of smooth, curved array heat transfer in the highly three-dimensional tunnel wall boundary layer flow and simulation of full-scale space shuttle vehicle pressure gradient levels was assessed.
Extending the absolute flux density scale at microwave wavelengths, the absolute flux densities at 22.285 GHz of several standard sources were determined using the absolute calibrations of the 6.1 meter antenna of the Hat Creek Observatory. Interpolation formulas for each nonthermal standard source have been derived by combining these data with those determined at lower frequencies. The suitability of employing the standard sources for calibrating other antennas is discussed.
A quantitative model is presented to account for the observed features of the binary pulsar PSR1953+53. A lower giant branch star is suggested to have evolved on a nuclear time scale while orbiting a neutron star, with mass transfer to the neutron star commencing when the giant attained its Roche lobe. A radius-core mass relationship was used to examined the processes that occurred after the hydrogen atmosphere was drained off to the neutron star. Tidal dissipation has circularized the orbit of the companion during mass transfer, while the orbit expanded. Accretion torques caused spin-up of the neutron star, which had a magnetic moment of no more than 3 x 10 to the 27th G/cu cm. The drained star became a degenerate dwarf with very faint luminosity, which accounts for no visual companion having been observed to date.
Some reddened quasars appear to be transitional objects in the paradigm of merger-induced black hole growth/ galaxy evolution, where a heavily obscured nucleus starts to be unveiled by powerful quasar winds evacuating the surrounding cocoon NuSTAR and XMM-Newton/Chandra observations of FIRST-2MASS-selected red quasars F2M 0830+3759 and F2M 1227+3214. We find that though F2M 0830 +3759 is moderately obscured N(sub H) = (2.1 +/- 0.2) x 10 (exp 22) per square centimeter) and F2M 1227+3214 is mildly absorbed (N(sub H),Z =3.4(+0.8/-0.7) X 10(exp -2) along the line of sight, heavier global obscuration may be present in both sources, with N(sub H) = 3.7 (+4.1/-2.6) X 10 (exp 23) per square centimeter) and less than 5.5 X 10(exp 23) per square centimeter) for F2M 0830+3759 and F2M 1227+ 3214, respectively. F2M 0830+3759 also has an excess of soft X-ray emission below 1 of dust and gas. Hard X-ray observations are able to peer through this gas and dust, revealing the properties of circumnuclear obscuration. Here, we present keV, which is well accommodated by a model where 7% of the intrinsic X-ray emission from the active galactic nucleus (AGN) is scattered into the line of sight. While F2M 1227+3214 has a dust-to-gas ratio (E(B - V)/NH) consistent with the Galactic value, the value of E(B - V)/NH for F2M 0830+3759 is lower than the Galactic standard, consistent with the paradigm that the dust resides on galactic scales while the X-ray reprocessing gas originates within the dust sublimation zone of the broad-line region. The X-ray and 6.1 μm luminosities of these red quasars are consistent with the empirical relations derived for high-luminosity, unobscured quasars, extending the parameter space of obscured AGNs previously observed by NuSTAR to higher luminosities.
A Digital Terrain Model (DTM) of Mars was derived with both Mercator and Sinusoidal Equal-Area projections from the global topographic map of Mars (scale 1:15 million, contour interval 1 km). Elevations on the map are referred to Mars' topographic datum that is defined by the gravity field at a 6.1-millibar pressure surface with respect to the center of mass of Mars. The DTM has a resolution at the equator of 1/59.226 degrees (exactly 1 km) per pixel. By using the DTM, the volumetric distribution of Mars topography above and below the datum has previously been calculated. Three types of elevation distributions of Mars' topography were calculated from the same DTM: (1) the frequency distribution of elevations at the pixel resolution; (2) average elevations in increments of 6 degrees in both longitude and latitude; and (3) average elevations in 36 separate blocks, each covering 30 degrees of latitude and 60 degrees of longitude.
A Digital Terrain Model (DTM) was derived with both Mercator and Sinusoidal Equal-area projections from the global topographic map of Mars at a scale of 1:15 million and a contour interval of 1 km. Elevations on the map are referred to the Mars topographic datum that is defined by the gravity field at a 6.1-millibar pressure surface with respect to the center of mass of Mars. The DTM has a resolution at the equator of 1/59.226 degrees (exactly 1 km) per pixel. By using the DTM, color-coded global maps of Mars' topography were generated in both the Mercator projection and the Sinusoidal Equal-Area projection. On both maps, colors indicate 1 km increments of height. From the equal-are dataset, the positive and negative elevation distributions are calculated to be 67 and 33 percent, respectively.
Tests were conducted in the NASA Langley 30 by 60 foot Wind Tunnel on a full scale 7.31 m (24 ft) long tip section of a wind turbine rotor blade. The blade tip section was built with ailerons on the trailing edge. The ailerons, which spanned a length of 6.1 m (20 ft), were designed so that two types could be evaluated: the plain and the balanced. The ailerons were hinged on the suction surface at the 0.62 X chord station behind the leading edge. The purpose of the tests was to measure the aerodynamic characteristics of the blade section for: an angle of attack range from 0 deg to 90 deg aileron deflections from 0 deg to -90 deg, and Reynolds numbers of 0.79 and 1.5 x 10 to the 6th power. These data were then used to determine which aileron configuration had the most desirable rotor control and aerodynamic braking characteristics. Tests were also run to determine the effects of vortex generators, leading edge roughness, and the gaps between the aileron sections on the lift, drag, and chordwise force coefficients of the blade tip section.
A spin tunnel study is reported on a scale model of a research airplane typical of low-wing, single-engine, light general aviation airplanes to determine the tail parachute diameter and canopy distance (riser length plus suspension-line length) required for energency spin recovery. Nine tail configurations were tested, resulting in a wide range of developed spin conditions, including steep spins and flat spins. The results indicate that the full-scale parachute diameter required for satisfactory recovery from the most critical conditions investigated is about 3.2 m and that the canopy distance, which was found to be critical for flat spins, should be between 4.6 and 6.1 m.
Composite materials are increasingly being considered and used for launch-vehicle structures. For shell structures, such as interstages, skirts, and shrouds, honeycomb-core sandwich composites are often selected for their structural efficiency. Therefore, it is becoming increasingly important to understand the structural response, including buckling, of sandwich composite shell structures. Additionally, small geometric imperfections can significantly influence the buckling response, including considerably reducing the buckling load, of shell structures. Thus, both the response of the theoretically perfect structure and the buckling imperfection sensitivity must be considered during the design of such structures. To address the latter, empirically derived design factors, called buckling knockdown factors (KDFs), were developed by NASA in the 1960s to account for this buckling imperfection sensitivity during design. However, most of the test-article designs used in the development of these recommendations are not relevant to modern launch-vehicle constructions and material systems, and in particular, no composite test articles were considered. Herein, a two-part study on composite sandwich shells to (1) examine the relationship between the buckling knockdown factor and the areal mass of optimized designs, and (2) to interrogate the imperfection sensitivity of those optimized designs is presented. Four structures from recent NASA launch-vehicle development activities are considered. First, designs optimized for both strength and stability were generated for each of these structures using design optimization software and a range of buckling knockdown factors; it was found that the designed areal masses varied by between 6.1% and 19.6% over knockdown factors ranging from 0.6 to 0.9. Next, the buckling imperfection sensitivity of the optimized designs is explored using nonlinear finite-element analysis and the as-measured shape of a large-scale composite cylindrical shell. When compared with the current buckling design recommendations, the results suggest that the current recommendations are overly conservative and that the development of new recommendations could reduce the acreage areal mass of many composite sandwich shell designs by between 4% and 19%, depending on the structure.
Topographic contour maps of Mars are compiled by synthesizing data acquired from various scientific experiments of the Mariner 9 mission, including S-band radio occultation, the UV spectrometer, the IR radiometer, the IR interferometer spectrometer, and television imagery, as well as earth-based radar information. The entire planet is mapped at scales of 1:25,000,000 and 1:5,000,000 using Mercator, Lambert, and polar stereographic map projections. For the computation of map projections, a biaxial spheroid figure is adopted. The semimajor and semiminor axes are 3393.4 and 3375.7 km, respectively, with a polar flattening of 0.0052. For the computation of elevations, a topographic datum is defined by a gravity field described in terms of spherical harmonics of fourth order and fourth degree combined with a 6.1-mbar occultation pressure surface. This areoid can be approximated by a triaxial ellipsoid with semimajor axes of A = 3394.6 km and B = 3393.3 km and a semiminor axis of C = 3376.3 km. The semimajor axis A intersects the Martian surface at longitude 105 deg W. The dynamic flattening of Mars is 0.00525. The contour interval of the maps is 1 km. For some prominent features where overlapping pictures from Mariner 9 are available, local contour maps at relatively larger scales were also compiled by photogrammetric methods on stereo plotters.