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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 163 records · Page 9
All sourcesNASA NTRSDOE OSTI
NASA NTRS2011

Sloshing of the Magnetized Cool Gas in the Cores of Galaxy Clusters

X-ray observations of many clusters of galaxies reveal the presence of edges in surface brightness and temperature, known as "cold fronts". In relaxed clusters with cool cores, these edges have been interpreted as evidence for the "sloshing" of the core gas in the cluster's gravitational potential. The smoothness of these edges has been interpreted as evidence for the stabilizing effect of magnetic fields "draped" around the front surfaces. To check this hypothesis, we perform high-resolution magnetohydrodynamics simulations of magnetized gas sloshing in galaxy clusters initiated by encounters with subclusters. We go beyond previous works on the simulation of cold fronts in a magnetized intracluster medium by simulating their formation in realistic, idealized mergers with high resolution ((Delta)x approx. 2 kpc). Our simulations sample a parameter space of plausible initial magnetic field strengths and field configurations. In the simulations, we observe strong velocity shears associated with the cold fronts amplifying the magnetic field along the cold front surfaces, increasing the magnetic field strength in these layers by up to an order of magnitude, and boosting the magnetic pressure up to near-equipartition with thermal pressure in some cases. In these layers, the magnetic field becomes strong enough to stabilize the cold fronts against Kelvin-Helmholtz instabilities, resulting in sharp, smooth fronts as those seen in observations of real clusters. These magnetic fields also result in strong suppression of mixing of high and low-entropy gas in the cluster, seen in our simulations of mergers in the absence of a magnetic field. As a result, the heating of the core due to sloshing is very modest and is unable to stave off a cooling catastrophe.

ZuHone, J. A.↗
NASA NTRS1957

Some Effects of Bluntness on Boundary-Layer Transition and Heat Transfer at Supersonic Speeds

Large downstream movements of transition observed when the leading edge of a hollow cylinder or a flat plate is slightly blunted are explained in terms of the reduction in Reynolds number at the outer edge of the boundary layer due to the detached shock wave. The magnitude of this reduction is computed for cones and wedges for Mach numbers to 20. Concurrent changes in outer-edge Mach number and temperature occur in the direction that would increase the stability of the laminar boundary layer. The hypothesis is made that transition Reynolds number is substantially unchanged when a sharp leading edge or tip is blunted. This hypothesis leads to the conclusion that the downstream movement of transition is inversely proportional to the ratio of surface Reynolds number with blunted tip or leading edge to surface Reynolds number with sharp tip or leading edge. The conclusion is in good agreement with the hollow-cylinder result at Mach 3.1.

Moeckel, W E↗
NASA NTRS1953

A study of the use of leading-edge notches as a means for improving the low-speed pitching-moment characteristics of a thin 45-degree swept wing of aspect ratio 4

In an effort to develop a simple but effective means for controlling the stall progression of swept wings, a low-speed investigation has been undertaken primarily to determine the effectiveness of leading-edge notches. The investigation included consideration of notch size, location, and shape, as well as combined effects of notch—chord-extension and notch-fence arrangements. The wing used in the study had 45° quarter-chord sweep, aspect ratio 4, taper ratio 0.3, and NACA 65A006 airfoil sections streamwise. The results of the investigation indicated that although substantial improvements were obtained in the wing pitching-moment characteristics at moderate angles of attack, by use of outboard leading-edge notches, none of these notches produced desirable wing stability characteristics throughout the angle-of-attack range. The wing pitching-moment characteristics were relatively insensitive to notch size and to small changes in span-wise location. The use of a notch at the inboard end of a chord-extension increased the effectiveness of the chord-extension near stall and produced the same results as a chord-extension having an equivalent inboard overhang. Small localized notch—chord-extension combinations were found to produce the same pitching-moment characteristics as much larger chord-extension configurations. A leading-edge root cutout used in conjunction with a small outboard notch—chord-extension combination resulted in fairly desirable overall wing pitching-moment characteristics.

Joseph Weil↗
NASA NTRS1955

Longitudinal Stability and Control Characteristics as Determined by the Rocket-Model Technique for an Inline, Cruciform, Canard Missile Configuration with a Low-Aspect-Ratio Wing Having Trailing-Edge Flap Controls for a Mach Number Range of 0.7 to 1.8

Two full-scale models of an inline, cruciform, canard missile configuration having a low-aspect-ratio wing equipped with flap-type controls were flight tested in order to determine the missile's longitudinal aerodynamic characteristics. Stability derivatives and control and drag characteristics are presented for a range of Mach number from 0.7 to 1.8. Nonlinear lift and moment curves were noted for the angle-of-attack range of this test (0 deg to 8 deg ). The aerodynamic-center location for angles of attack near 5 deg remained nearly constant for supersonic speeds at 13.5 percent of the mean aerodynamic chord; whereas for angles of attack near O deg, there was a rapid forward movement of the aerodynamic center as the Mach number increased. At a control deflection of O deg, the missile's response to the longitudinal control was in an essentially fixed space plane which was not coincident with the pitch plane as a result of the missile rolling. As a consequence, stability characteristics were determined from the resultant of pitch and yaw motions. The damping-in-pitch derivatives for the two angle-of-attack ranges of the test are in close agreement and varied only slightly with Mach number. The horn-balanced trailing-edge flap was effective in producing angle of attack over the Mach number range.

Baber, H. T., Jr.↗
NASA NTRS1955

Longitudinal Stability and Control Characteristics as Determined by the Rocket-model Technique for an Inline, Cruciform, Canard Missile Configuration with a Low-aspect-ratio Wing Having Trailing-edge Flap Controls for a Mach Number Range of 0.7 to 1.

Two full-scale models of an inline, cruciform, canard missile configuration having a low-aspect-ratio wing equipped with flap-type controls were flight tested in order to determine the missile's longitudinal aerodynamic characteristics. Stability derivatives and control and drag characteristics are presented for a range of Mach number from 0.7 to 1.8. Nonlinear lift and moment curves were noted for the angle - of-attack range of this test (0 deg to 8 deg). The aerodynamic-center location for angles of attack near 50 remained nearly constant for supersonic speeds at 13.5 percent of the mean aerodynamic chord; whereas for angles of attack near 0 deg, there was a rapid forward movement of the aerodynamic center as the Mach number increased. At a control deflection of 0 deg, the missile's response to the longitudinal control was in an essentially fixed space plane which was not coincident with the pitch plane as a result of the missile rolling. As a consequence, stability characteristics were determined from the resultant of pitch and yaw motions. The damping-in-pitch derivatives for the two angle -of-attack ranges of the test are in close agreement and varied only slightly with Mach number. The horn-balanced trailing-edge flap was effective in producing angle of attack over the Mach number range.

Baber, Hal T , Jr↗
NASA NTRS1987

Wind-tunnel investigation of a full-scale general aviation airplane equipped with an advanced natural laminar flow wing

An investigation was conducted in the Langley 30- by 60-Foot Wind Tunnel to evaluate the performance, stability, and control characteristics of a full-scale general aviation airplane equipped with an advanced laminar flow wing. The study focused on the effects of natural laminar flow and advanced boundary layer transition on performance, stability, and control, and also on the effects of several wing leading edge modifications on the stall/departure resistance of the configuration. Data were measured over an angle-of-attack range from -6 to 40 deg and an angle-of-sideslip range from -6 to 20 deg. The Reynolds number was varied from 1.4 to 2.4 x 10 to the 6th power based on the mean aerodynamic chord. Additional measurements were made using hot-film and sublimating chemical techniques to determine the condition of the wing boundary layer, and wool tufts were used to study the wing stall characteristics. The investigation showed that large regions of natural laminar flow existed on the wing which would significantly enhance cruise performance. Also, because of the characteristics of the airfoil section, artificially tripping the wing boundary layer to a turbulent condition did not significantly effect the lift, stability, and control characteristics. The addition of a leading-edge droop arrangement was found to increase the stall angle of attack at the wingtips and, therefore, was considered to be effective in improving the stall/departure resistance of the configuration. Also the addition of the droop arrangement resulted in only minor increases in drag.

Murri, Daniel G.↗
NASA NTRS1975

Low-speed wind-tunnel tests of a one-tenth-scale model of a blended-arrow advanced supersonic transport

Tests were conducted in the Langley full scale tunnel to determine the low-speed aerodynamic characteristics of a 1/10 scale model of a blended-arrow advanced supersonic transport. Tests were made for the clean configuration and a high-lift configuration with several combinations of leading- and trailing-edge flaps deflected for providing improved lift and longitudinal stability in the landing and takeoff modes. The tests were conducted for a range of angles of attack from about -6 deg to 30 deg, sideslip angles from -5 deg to 10 deg, and for Reynolds numbers from 6.78 x 1,000,000 to 13.85 x 1,000,000 corresponding to test velocities of 41 knots to 85 knots, respectively.

Lemore, H. C.↗
NASA NTRS1984

Pattern selection in solidification

Directional solidification of alloys produces a wide variety of cellular or lamellar structures which, depending upon growth conditions, may be reproducibly regular or may behave chaotically. It is not well understood how these patterns are selected and controlled or even whether there ever exist sharp selection mechanisms. A related phenomenon is the spatial propagation of a pattern into a system which has been caused to become unstable against pattern-forming deformations. This phenomenon has some features in common with the propagation of sidebranching modes in dendritic solidification. In a class of one-dimensional models, the nonlinear system can be shown to select the propagating mode in which the leading edge of the pattern is just marginally stable. This stability principle, when applicable, predicts both the speed of propagation and the geometrical characteristics of the pattern which forms behind the moving front. A boundary-layer model for fully two or three dimensional solidification problems appears to exhibit similar mathematical behavior.

Langer, J. S.↗
NASA NTRS2004

Fundamental Investigations of Airframe Noise

An extensive numerical and experimental study of airframe noise mechanisms associated with a subsonic high-lift system has been performed at NASA Langley Research Center (LaRC). Investigations involving both steady and unsteady computations and experiments on a small-scale, part-span flap model are presented. Both surface (steady and unsteady pressure measurements, hot films, oil flows, pressure sensitive paint) and off surface (5 hole-probe, particle-imaged velocimetry, laser velocimetry, laser light sheet measurements) were taken in the LaRC Quiet Flow Facility (QFF) and several hard-wall tunnels up to flight Reynolds number. Successful microphone array measurements were also taken providing both acoustic source maps on the model, and quantitative spectra. Critical directivity measurements were obtained in the QFF. NASA Langley unstructured and structured Reynolds- Averaged Navier-Stokes codes modeled the flap geometries excellent comparisons with surface and offsurface experimental data were obtained. Subsequently, these meanflow calculations were utilized in both linear stability and direct numerical simulations of the flap-edge flow field to calculate unsteady surface pressures and farfield acoustic spectra. Accurate calculations were critical in obtaining not only noise source characteristics, but shear layer correction data as well. Techniques utilized in these investigations as well as brief overviews of results will be given.

Macaraeg, M. G.↗
NASA NTRS2015

Synthesis and Modification of Holey Graphene for Energy Storage

Porous graphene materials have recently drawn significant interest for energy storage applications due to the effective use of the graphitic surface area provided by the pores. Holey graphene is a type of porous graphene with pores in the form of physical through thickness holes. This type of porous graphene provides a unique morphology that allows for effective transport of molecular species through the graphene plane. Therefore, as we recently demonstrated, holey graphene could be advantageous in the preparation of electrodes in supercapacitors with high volumetric performance.1 We have developed methodologies to scalably prepare holey graphene materials with or without the use of catalysts (Figure 1).1,2 In this presentation, we will discuss the choice of synthetic strategies and experimental conditions on the morphology and physical properties of holey graphene (Figure 2) and the effect on their electrochemical properties. We will also present the results on the further modification of holey graphene to study their physical and chemical properties, especially with regard to the stability and chemical reactivity of carbons located around the edges of the holes.

Lin, Yi↗
NASA NTRSDate not supplied

Design and Characterization of New 90 GHz Detectors for the Cosmology Large Angular Scale Surveyor (CLASS)

The Cosmology Large Angular Scale Surveyor (CLASS) is a polarization-sensitive telescope array located at an altitude of 5,200 m in the Chilean Atacama Desert. CLASS is designed to measure “E-mode” (even parity) and “B-mode” (odd parity) polarization patterns in the Cosmic Microwave Background (CMB) over large angular scales with the aim of improving our understanding of inflation, reionization, and dark matter. CLASS is currently observing with three telescopes covering four frequency bands: one at 40 GHz (Q); one at 90 GHz (W1); and one dichroic system at 150/220 GHz (G). In these proceedings, we discuss the updated design and in-lab characterization of new 90 GHz detectors. The new detectors include design changes to the transition-edge sensor (TES) bolometer architecture, which aim to improve stability and optical efficiency. We assembled and tested four new detector wafers, to replace four modules of the W1 focal plane. These detectors were installed into the W1 telescope, and will achieve first light in the austral winter of 2022. We present electrothermal parameters and bandpass measurements from in-lab dark and optical testing. From in-lab dark tests, we also measure a median NEP of 12.3 aW √ s across all four wafers about the CLASS signal band, which is below the expected photon NEP of 32 aW √ s from the field. We therefore expect the new detectors to be photon noise limited."

Carolina Nunez↗
NASA NTRS1991

Application of linear stability theory in laminar flow design

The linear stability of fully three-dimensional supersonic boundary layers formed over swept-wing configurations is investigated using a modified version of the linear stability code COSAL. Configurations studied include a highly swept leading-edge model to be utilized for transition studies in the LARC Low-disturbance Mach 3.5 Pilot Tunnel. The model is a representation of the leading edge of a laminar flow control wing for the F-16XL aircraft. In addition, the region over a laminar flow control glove fitted on the midportion of an F-16XL wing was studied. For each configuration, estimates of the location of the onset of transition were computed using linear stability theory and the e exp N method. The effectiveness of suction in stabilizing the boundary layer over the F-16XL wing glove was also investigated.

Iyer, Venkit↗
NASA NTRS1987

Plasma parameters in the near wake of the Space Shuttle

Results are presented from a Spherical Retarding Potential Analyzer and a spherical Langmuir Probe which were flown on the Spacelab-2 mission of the Space Shuttle Orbiter in August 1985. The probes were mounted in such a position that for certain Orbiter orientations, the wake edge formed by the starboard payload bay sill swept across the probes at a roll rate of once per orbit about an inertially stabilized pointing direction. Profiles of O(+) number density and electron number density across the wake edge showed that the charge particle number densities declined into the wake at a faster rate than that predicted by theoretical models. At angles of more than 10 deg into the wake from its geometric edge, two electron populations are found, with different mean random energies and electrical potentials. One of these populations is suggested to be the result of selective filtering of the high energy tails of the ambient ionospheric population, the other, a result of photoemission from the payload bay surfaces.

Raitt, W. J.↗
NASA NTRS1945

Effects of Wing and Nacelle Modifications on Drag and Wake Characteristics of a Bomber-Type Airplane Model

An investigation of a model of a large four-engine bomber was conducted in the Langley 19-f'oot pressure tunnel to determine the effects of several wing and nacelle modifications on drag characteristics and air­ flow characteristics at the tail. Leading-edge gloves, trailing-edge extensions, and modified nacelle afterbodies were tested individual ly and in combination. The effects of the various modifications were determined by force tests, tuft observations, and turbulence s1ITveys in the region of the tail. Tests were made with fixed and natural transition on the wing and with propellers operating and propellers off. Most of the tests were con- ducted at a Reynolds number of approximately 2.6 x 106. The results indicated that application of certain of the modifications provided worth-while improvements in the characteristics or the model. The flow over the wing and flaps was improved, the drag was reduced, and the turbulence in the region of the tail was reduced. Trailing-edge extensions were the most effective individual modification in improving the flow over the wing with wing flaps neutral, cowl and intercooler flaps clos ed. Modified nacelle afterbodies were the most effectiv8 individual edification in reducing drag with either fixed or natural transition on the wing; however, trailin6-edge extensions were slightly more effective with fixed transition. Combinations of either leading­ or trailing-edge extensions and modified afterbodies were more effective than either modification alone. With cowl and intercooler flaps open, trailing-edge extensions with modified afterbodies provided substantial improvement in flow and drag characteristics. With wing flaps deflected, enclosing the flap behind the inboard nacelle within an extended afterbody or cutting the flaps at the nacelle appeared. to be the most promising methods of improving the f low over the flaps and the tail. Although the results of hot-wire-anenometer surveys were not conclusive in regard to buffeting characteristics, the modifications did educe the turbulence at the tail with wing flaps both neutral and deflected. The modifications, as a rule, were favorable to maximum lift. Appreciable reductions in longitudinal stability of the model were caused by addition of leading -edge gloves and tr ailing -edge extensions.

Neely, Robert H.↗
NASA NTRS1949

The Effects of Horizontal-Tail Location and Wing Modifications on the High-Speed Stability and Control Characteristics of a 01.17-Scale Model of the McDonnell XF2H-1 Airplane (TED No, NACA DE336)

An additional series of high-speed wind-tunnel tests of a modified 0.17-scale model of the McDonnell XF2H-1 airplane was conducted to evaluate the effects of a reduction in the thickness-to-chord ratios of the tail planes, the displacement of the horizontal tail relative to the vertical tail, and the extension of the trailing edge of the wing. Two tail-intersection fairings designed to improve the flow at the tail were also tested. The pitching-moment characteristics of the model were improved slightly by the use of the thinner tail sections. Rearward or rearward and downward displacements of the horizontal tail increased the critical Mach number at the tail intersection from 0.725 to a maximum of 0.80, but caused an excessive change in pitching-moment coefficient at the higher Mach numbers. Extending the trailing edge of the wing did not improve the static longitudinal-stability characteristics, but increased the pitching-down tendency between 0.725 and 0.825 Mach numbers prior to the pitching-up tendency. The extended wing did, however, increase the Mach numbers at which these tendencies occurred. The increase in the Mach numbers of divergence and the tuft studies indicate a probable increase in the buffet limit of the prototype airplane. No perceptible improvement of flow at the tail intersection was observed with the two fairings tested on the forward tail configuration.

Emerson, Horace F.↗
NASA NTRS1987

High-angle-of-attack stability and control improvements for the EA-6B Prowler

The factors involved in high-angle-of-attack directional divergence phenomena for the EA-6B ECM aircraft have been investigated in NASA-Langley wind tunnel facilities in order to evaluate airframe modifications which would eliminate or delay such divergence to angles-of-attack farther removed from the operational flight envelope of the aircraft. The results obtained indicate that an adverse sidewash at the aft fuselage and vertical tail location is responsible for the directional stability loss, and that the sidewash is due to a vortex system generated by the fuselage-wing juncture. Modifications encompassing a wing inboard leading edge droop, a wing glove strake, and a vertical fin extension, have significantly alleviated the stability problem.

Jordan, Frank L.↗