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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 469 records · Page 26
All sourcesNASA NTRSDOE OSTI
NASA NTRS1992

Growth and characterization of lead bromide crystals

Lead(II) bromide was purified by a combination of directional freezing and zone-refining methods. Differential thermal analysis of the lead bromide showed that a destructive phase transformation occurs below the melting temperature. This transformation causes extensive cracking, making it very difficult to grow a large single crystal. Energy of phase transformation for pure lead bromide was determined to be 24.67 cal/g. To circumvent this limitation, crystals were doped by silver bromide which decreased the energy of phase transformation. The addition of silver helped in achieving the size, but enhanced the inhomogeneity in the crystal. The acoustic attenuation constant was almost identical for the pure and doped (below 3000 ppm) crystals.

Singh, N. B.↗
NASA NTRS1993

Application of a Flush Airdata Sensing System to a Wing Leading Edge (LE-FADS)

The feasibility of locating a flush airdata sensing (FADS) system on a wing leading edge where the operation of the avionics or fire control radar system will not be hindered is investigated. The leading-edge FADS system (LE-FADS) was installed on an unswept symmetrical airfoil and a series of low-speed wind-tunnel tests were conducted to evaluate the performance of the system. As a result of the tests it is concluded that the aerodynamic models formulated for use on aircraft nosetips are directly applicable to wing leading edges and that the calibration process is similar. Furthermore, the agreement between the airdata calculations for angle of attack and total pressure from the LE-FADS and known wind-tunnel values suggest that wing-based flush airdata systems can be calibrated to a high degree of accuracy. Static wind-tunnel tests for angles of attack from -50 deg to 50 deg and dynamic pressures from 3.6 to 11.4 lb/sq ft were performed.

Whitmore, Stephen A.↗
NASA NTRS1992

Flight test operations using an F-106B research airplane modified with a wing leading-edge vortex flap

Flight tests of an F-106B aircraft equipped with a leading-edge vortex flap, which represented the culmination of a research effort to examine the effectiveness of the flap, were conducted at the NASA Langley Research Center. The purpose of the flight tests was to establish a data base on the use of a wing leading-edge vortex flap as a means to validate the design and analysis methods associated with the development of such a vortical flow-control concept. The overall experiment included: refinements of the design codes for vortex flaps; numerous wind tunnel entries to aid in verifying design codes and determining basic aerodynamic characteristics; design and fabrication of the flaps, structural modifications to the wing tip and leading edges of the test aircraft; development and installation of an aircraft research instrumentation system, including wing and flap surface pressure measurements and selected structural loads measurements; ground-based simulation to assess flying qualities; and finally, flight testing. This paper reviews the operational aspects associated with the flight experiment, which includes a description of modifications to the research airplane, the overall flight test procedures, and problems encountered. Selected research results are also presented to illustrate the accomplishments of the research effort.

Dicarlo, Daniel J.↗
NASA NTRS1993

Effect of leading-edge porosity on blade-vortex interaction noise

The effect of the porous leading-edge of an airfoil on the blade-vortex interaction noise, which dominates far-field acoustic spectrum of the helicopter, is investigated. The thin-layer Navier-Stokes equations are solved with a high-order upwind-biased scheme and a multizonal grid system. The Baldwin-Lomax turbulence model is modified for considering transpiration on the surface. The amplitudes of the propagating acoustic wave in the near-field are calculated directly from the computation. The porosity effect on the surface is modeled. Results show leading-edge transpiration can suppress pressure fluctuations at the leading-edge during BVI, and consequently reduce the amplitude of propagating noise by 30 percent at maximum in the near-field. The effect of porosity factor on the noise level is also investigated.

Lee, Soogab↗
NASA NTRS1993

Application of a flush airdata sensing system to a wing leading edge (LE-FADS)

This paper investigates the feasibility of locating a flush air-data sensing (FADS) system on a wing leading edge where the operation of the avionics or fire control radar system will not be hindered. The leading-edge FADS system (LE-FADS) was installed on an unswept symmetrical airfoil, and a series of low-speed wind-tunnel tests were conducted to evaluate the performance of the system. As a result of the tests it is concluded that the aerodynamic models formulated for use on aircraft nosetips are directly applicable to wing leading edges and that the calibration process is similar. Furthermore, the agreement between the air-data calculations for angle of attack and total pressure from the LE-FADS and known wind-tunnel values suggest that wing-based flush air-data systems can be calibrated to a high degree of accuracy. Static wind-tunnel tests for angles of attack from -50 to 50 deg and dynamic pressures from 3.6 to 11.4 lb/sq ft were performed.

Whitmore, Stephen A.↗
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 NTRS1956

Heat-requirements for Ice Protection of a Cyclically Gas-heated, 36 Degree Swept Airfoil with Partial-span Leading-edge Slat

Heating requirements for satisfactory cyclic de-icing over a wide range of icing and operating conditions have been determined for a gas-heated, 36deg swept airfoil of 6.9-foot chord with a partial-span leading-edge slat. Comparisons of heating requirements and effectiveness were made between the slatted and unslatted portions of the airfoil. Studies were also made comparing cyclic de-icing with continuous anti-icing, and cycll.cde-icing systems with and without leading-edge ice-free parting strips. De-icing heat requirements were approximately the same with either heated or unheated parting strips because of the aerodynamic effects of the 36deg sweep angle and the spanwise saw-tooth profile of leading-edge glaze-ice deposits. Cyclic de-icing heat-source requirements were found to be one-fourth or less of the heat requirements for complete anti-icing. The primary factors that affected the performance of the cyclic de-icing heating system were ambient air temperature, heat distribution, and thermal lag.

Gray, Vernon H↗
NASA NTRS1993

Lead poisoning: An overview

A problem that should be of great concern to all of us is the lead poisoning of children. First, I would like to present a short overview concerning the reasons everyone should care about lead poisoning, then discuss the history of lead poisoning, what is happening today across the country, and the future.

Gendel, Neil↗
NASA NTRS1992

Control of leading-edge vortices on a delta wing

The unsteady flow structure of leading-edge vortices on a delta wing has been investigated using new types of experimental techniques, in order to provide insight into the consequences of various forms of active control. These investigations involve global control of the entire wing and local control applied at crucial locations on or adjacent to the wing. Transient control having long and short time-scales, relative to the convective time-scale C/U(sub infinity), allows substantial modification of the unsteady and time-mean flow structure. Global control at long time-scale involves pitching the wing at rates an order of magnitude lower than the convective time-scale C/U(sub infinity), but at large amplitudes. The functional form of the pitching maneuver exerts a predominant influence on the trajectory of the feeding sheet, the instantaneous vorticity distribution, and the instantaneous location of vortex breakdown. Global control at short time-scales of the order of the inherent frequency of the shear layer separating from the leading-edge and the natural frequency of vortex breakdown shows that 'resonant' response of the excited shear layer-vortex breakdown system is attainable. The spectral content of the induced disturbance is preserved not only across the entire core of the vortex, but also along the axis of the vortex into the region of vortex breakdown. This unsteady modification results in time-mean alteration of the axial and swirl velocity fields and the location of vortex breakdown. Localized control at long and short time-scales involves application of various transient forms of suction and blowing using small probes upstream and downstream of the location of vortex breakdown, as well as distributed suction and blowing along the leading-edge of the wing applied in a direction tangential to the feeding sheet. These local control techniques can result in substantial alteration of the location of vortex breakdown; in some cases, it is possible to accomplish this without net mass addition to the flow field.

Magness, C.↗
NASA NTRS1994

Reduction of blade-vortex interaction noise through porous leading edge

The effect of the porous leading edge of an airfoil on the blade-vortex interaction noise, which dominates the far-field acoustic spectrum of the helicopter, is investigated. The thin-layer Navier-Stokes equations are solved with a high-order upwind-biased scheme and a multizonal grid system. The Baldwin-Lomax turbulence model is modified for considering transpiration on the surface. The amplitudes of the propagating acoustic wave in the near field are calculated directly from the computation. The porosity effect on the surface is modeled in two ways: (1) imposition of prescribed transpiration velocity distribution and (2) calculation of transpiration velocity distribution by Darcy's law. Results show leading-edge transpiration can suppress pressure fluctuations at the leading edge during blade-vortex interaction and consequently reduce the amplitude of propagating noise by 30% at a maximum in the near field.

Lee, Soogab↗
NASA NTRS1996

Experimental Surface Pressure Data Obtained on 65 deg Delta Wing Across Reynolds Number and Mach Number Ranges. Vol. 3: Medium-radius leading edge

An experimental wind tunnel test of a 65 deg delta wing model with interchangeable leading edges was conducted in the Langley National Transonic Facility (NTF). The objective was to investigate the effects of Reynolds and Mach numbers on slender-wing leading-edge vortex flows with four values of wing leading-edge bluntness. Experimentally obtained pressure data are presented without analysis in tabulated and graphical formats across a Reynolds number range of 6 x 10(exp 6) to 120 x 10(exp 6) at a Mach number of 0.85 and across a Mach number range of 0.4 to 0.9 at Reynolds numbers of 6 x 10(exp 6), 60 x 10(exp 6), and 120 x 10(exp 6). Normal-force and pitching-moment coefficient plots for these Reynolds number and Mach number ranges are also presented.

Chu, Julio↗
NASA NTRS1996

Experimental Surface Pressure Data Obtained on 65 deg Delta Wing Across Reynolds Number and Mach Number Ranges. Vol. 4: Large-radius leading edge

An experimental wind tunnel test of a 65 deg delta wing model with interchangeable leading edges was conducted in the Langley National Transonic Facility (NTF). The objective was to investigate the effects of Reynolds and Mach numbers on slender-wing leading-edge vortex flows with four values of wing leading-edge bluntness. Experimentally obtained pressure data are presented without analysis in tabulated and graphical formats across a Reynolds number range of 6 x 10(exp 6) to 120 x 10(exp 6) at a Mach number of 0.85 and across a Mach number range of 0.4 to 0.9 at Reynolds numbers of 6 x 10(exp 6) and 60 x 10(exp 6). Normal-force and pitching-moment coefficient plots for these Reynolds number and Mach number ranges are also presented.

Chu, Julio↗
NASA NTRS1996

Improved Method for Prediction of Attainable Wing Leading-Edge Thrust

Prediction of the loss of wing leading-edge thrust and the accompanying increase in drag due to lift, when flow is not completely attached, presents a difficult but commonly encountered problem. A method (called the previous method) for the prediction of attainable leading-edge thrust and the resultant effect on airplane aerodynamic performance has been in use for more than a decade. Recently, the method has been revised to enhance its applicability to current airplane design and evaluation problems. The improved method (called the present method) provides for a greater range of airfoil shapes from very sharp to very blunt leading edges. It is also based on a wider range of Reynolds numbers than was available for the previous method. The present method, when employed in computer codes for aerodynamic analysis, generally results in improved correlation with experimental wing-body axial-force data and provides reasonable estimates of the measured drag.

Carlson, Harry W.↗
NASA NTRS1996

Experimental Surface Pressure Data Obtained on 65 deg Delta Wing Across Reynolds Number and Mach Number Ranges: Sharp Leading Edge - Volume 1

An experimental wind tunnel test of a 65 deg delta wing model with interchangeable leading edges was conducted in the Langley National Transonic Facility (NTF). The objective was to investigate the effects of Reynolds and Mach numbers on slender-wing leading-edge vortex flows with four values of wing leading-edge bluntness. Experimentally obtained pressure data are presented without analysis in tabulated and graphical formats across a Reynolds number range of 6 x 10(exp 6) to 36 x 10(exp 6) at a Mach number of 0.85 and across a Mach number range of 0.4 to 0.9 at a Reynolds number of 6 x 10(exp 6). Normal-force and pitching-moment coefficient plots for these Reynolds number and Mach number ranges are also presented.

Chu, Julio↗
NASA NTRS1996

Experimental Surface Pressure Data Obtained on 65 deg Delta Wing Across Reynolds Number and Mach Number Ranges: Small-Radius Leading Edge - Volume 2

An experimental wind tunnel test of a 65 deg. delta wing model with interchangeable leading edges was conducted in the Langley National Transonic Facility (NTF). The objective was to investigate the effects of Reynolds and Mach numbers on slender-wing leading-edge vortex flows with four values of wing leading-edge bluntness. Experimentally obtained pressure data are presented without analysis in tabulated and graphical formats across a Reynolds number range of 6 x 10(exp 6) to 84 x 10(exp 6) at a Mach number of 0.85 and across a Mach number range of 0.4 to 0.9 at Reynolds numbers of 6 x 10(exp 6) and 60 x 10(exp 6). Normal-force and pitching-moment coefficient plots for these Reynolds number and Mach number ranges are also presented.

Chu, Julio↗
NASA NTRS1960

Large-Scale Wind-Tunnel Tests and Evaluation of the Low-Speed Performance of a 35 deg Sweptback Wing Jet Transport Model Equipped with a Blowing Boundary-Layer-Control Flap and Leading-Edge Slat

A wind-tunnel investigation was conducted to determine the effect of trailing-edge flaps with blowing-type boundary-layer control and leading-edge slats on the low-speed performance of a large-scale jet transport model with four engines and a 35 deg. sweptback wing of aspect ratio 7. Two spanwise extents and several deflections of the trailing-edge flap were tested. Results were obtained with a normal leading-edge and with full-span leading-edge slats. Three-component longitudinal force and moment data and boundary-layer-control flow requirements are presented. The test results are analyzed in terms of possible improvements in low-speed performance. The effect on performance of the source of boundary-layer-control air flow is considered in the analysis.

Hickey, David H.↗
NASA NTRS1959

Transonic Flutter Characteristics of a 45 deg Sweptback Wing with Various Distributions of Ballast Along the Leading Edge

An investigation of the use of ballast at the leading edge of a sweptback wing as a flutter fix has been made. The investigation was conducted in the Langley transonic blowdown tunnel with wing models which had an aspect ratio of 4, sweepback of the quarter-chord line of 450, and a taper ratio of 0.2. Four ballast configurations, which included different amounts of ballast distributed at two different span-wise locations, were investigated. Full-span sting-mounted models were employed. Data were obtained over a Mach number range from 0.65 to 1.32. Comparison of the data for the ballasted wings with data for a similar wing without ballast shows that in the often critical Mach number range between 0.85 and 1.05, the dynamic pressure required for flutter is increased by as much as 100 percent due to the addition of about 6 percent of the wing mass as ballast at the leading edge of the outboard sections. Furthermore, there are indications that similar benefits of leading-edge ballast can be obtained at Mach numbers above M = 1.1. Changing the spanwise location of the ballast and increasing the amount of the ballast by a factor of about 2 had very little additional effect on the dynamic pressure required for flutter. The possibility, therefore, exists that the beneficial effects obtained may be accomplished by using less than the minimum of about 6 percent of the wing mass as ballast as investigated in this paper.

John R Unangst↗