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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 NTRS1986

Structureborne noise in aircraft: Modal tests

As part of an investigation to develop measurement techniques for structureborne noise, three modal surveys have been conducted on an OV-10A aircraft and the results have been presented. The purpose of the modal surveys was to identify suitable locations for mounting accelerometer and strain gages in subsequent tests in which transfer functions relating wing vibration to interior noise were to be determined. These surveys are as follows:(1) wing/fuselage modal survey utilizing one shaker under the right wing; (2) complete wing modal survey utilizing two shakers, one under each wing; and (3) fuselage side panel modal survey utilizing a small instrumented hammer. The predominant frequencies and damping ratios for each analysis were listed in tables. The primary mode shapes at the lower frequencies and at frequencies near the expected engine driving frequencies have been shown for each survey.

Clevenson, S. A.↗
NASA NTRS1934

The torsional stiffness of thin duralumin shells subjected to large torques

This report gives a simple method of estimating the torsional stiffness of thin shells, such as box beams or stressed-skin wings under large torque loads. A general efficiency chart for shells in torsion is established, based on the assumption that the efficiency of the web sheet in resisting deformation decreases linearly with the average stress. The chart is used to calculate the torsional deflections of eight box beams, a test wing panel, and a complete wing; the results of the calculations are shown in comparison with the test results. The agreement is probably as good as might be expected considering the empirical nature of the method and the well-known dispersion between results of tests on thin sheet-metal structures.

Kuhn, Paul↗
NASA NTRS1958

The Prospects for Laminar Flow on Hypersonic Airplanes

The factors which affect the extent of laminar flow on airplanes for hypersonic flight are discussed on the basis of the available data. Factors considered include flight Reynolds number, surface roughness, angle of attack, angle of leading-edge sweepback, and aerodynamic interference. Test data are presented for one complete configuration.

MACH NUMBER EFFECTS - COMPLETE WINGS↗
NASA NTRS1989

A new airfoil design concept

The present airfoil design concept is based on utilizing unconventional geometry characteristics near the airfoil trailing edge which include a finite trailing edge thickness, strongly divergent trailing edge upper and lower surfaces, and high surface curvature on the lower surface at or near the lower surface trailing edge. This paper presents computational analyses of airfoils and a wing utilizing the concept, airfoil validation wind tunnel test results of several configurations, and wing-validation wind tunnel test results for a complete wing design. In addition to validating the concept, the airfoil and wing testing provided additional detailed data to better understand the aerodynamic advantage of such an unconventional trailing edge configuration. It is demonstrated that the concept represents a significant step in airfoil technology beyond that achieved with the Supercritical Airfoil. This concept provides the aerodynamicist an additional degree of design freedom and flexibility previously unrecognized.

Henne, P. A.↗
NASA NTRS1934

The calculation of lateral stability with free controls

The discussion of the structural methods for obtaining lateral stability discloses the remarkable influence of the constant fuselage and wing proportions to the yawing moments. For the effectiveness of modifications in vertical tail surfaces and tail length, these quotas - little observed heretofore, in this connection - are decisive. This also applies to the amount of dihedral of the wing with regard to the roll stability of the complete wing already existing without angle of the dihedral.

Mathias, Gotthold↗
NASA NTRS1941

Wind-Tunnel Investigation of an NACA Low-Drag Tapered Wing with Straight Trailing Edge and Simple Split Flaps, Special Report

An investigation was conducted in the NACA 19-foot pressure wind tunnel of a tapered wing with straight railing edge having NACA 66 series low-drag airfoil sections and equipped with full-span and partial-span simple split flaps. The airfoil sections used were the NACA 66,2-116 at the root and the 66,2-216 at the tip. The primary purpose of the investigation was to determine the effect of the split flaps on the aerodynamic characteristics of the tapered wing. Complete lift, drag, and pitching-moment coefficients were determined for the plain wing and for each flap arrangement through a Reynold number range of 2,600,000 to 4,600,000. The results of this investigation indicate that values of maximum lift coefficient comparable to values obtained on tapered wings with conventional sections and similar flap installations can be obtained from wings with the NACA low-drag sections. The increment of maximum lift due to the split flap was found to vary somewhat with Reynold number over the range investigated. The C(sub L)max of the wing alone is 1.49 at a Reynolds number of 4,600,000; whereas with the partial-span simple split flap it is 2.22 and with the full-span arrangement, 2.80. Observations of wool tufts on the wing indicate that the addition of split flaps did not appreciable alter the pattern of the stall; even though the stall did occur more abruptly than with the wing alone.

Muse, Thomas C.↗
NASA NTRS1996

Simulation of Three-Dimensional Symmetric and Asymmetric Instabilities in Attachment-Line Boundary Layers

On a swept wing, contamination along the leading edge, Tollmien-Schlichting waves, stationary or traveling crossflow vortices, and/or Taylor-Gortler vortices can cause the catastrophic breakdown of laminar to turbulent flow, which leads to increased skin-friction drag for the aircraft. The discussion in this Note will be limited to disturbances which evolve along the attachment line (leading edge of swept wing). If the Reynolds number of the attachment-line boundary layer is greater than some critical value, then the complete wing is inevitably engulfed in turbulent flow. Essentially, there are two critical Reynolds number points that must be considered. The first is for small-amplitude disturbances, and the second is for bypass transition. The present study will use direct numerical simulations to validate a linear 2D-eigenvalue prediction method based on parabolized stability equations by Lin and Malik. This method is considered because it suggests that a number of symmetric and asymmetric modes exist and are stable or unstable on the attachment line depending on the Reynolds number. If validated, the approach would predict a number of modes which are linearly damped in the Reynolds number regime 100 to 245; however, these modes may grow nonlinearly and provide an explanation to this region.

Joslin, Ronald D.↗
NASA NTRS1941

Wind-Tunnel Investigation of an NACA 66,2-216 Low-Drag Wing with Split Flaps of Various Sizes, Special Report

An investigation was conducted in the NACA 19-foot pressure wind tunnel of a rectangular wing having NACA 66, 2-216 low-drag airfoil sections and various sizes of simple split flaps. The purpose of the investigation was, primarily, to determine the influence of these flap installations on the aerodynamic characteristics of the wing. Complete lift, drag, and pitching-moment characteristics were determined for a range of test Reynolds numbers from about 2,600,000 to 4,600,000 for each of the installations and for the plain wing. The results of this investigation indicate that values of maximum lift coefficient similar to those of wings with conventional airfoil sections and split flaps can be expected of wings having the NACA 66,2-216 low-drag sections. The increment of maximum lift due to the split flap was found to be practically independent of the Reynolds number over the range investigated. The optimum split flap on the basis of maximum lift appears to have a chord about 20% of the wing chord and a deflection of 60 degrees. The C(sub L) max of the wing with the 0.20c partial-span flap deflected 60 degrees is 2.07 at a Reynolds number of 4,600,000 while with the full-span flap it is approximately 2.53; the increment of the maximum lift coefficient due to the flap is approximately proportional to the flap span. Although the addition of a split flap tends to hasten the stall and to cause it to occur more abruptly, little change in pattern is evidenced by observations of the behavior of wool tufts on the wing.

Muse, Thomas C.↗
NASA NTRS1999

Pressure-Sensitive Paint and Video Model Deformation Systems at the NASA Langley Unitary Plan Wind Tunnel

Pressure-sensitive paint (PSP) and video model deformation (VMD) systems have been installed in the Unitary Plan Wind Tunnel at the NASA Langley Research Center to support the supersonic wind tunnel testing requirements of the High Speed Research (HSR) program. The PSP and VMD systems have been operational since early 1996 and provide the capabilities of measuring global surface static pressures and wing local twist angles and deflections (bending). These techniques have been successfully applied to several HSR wind tunnel models for wide ranges of the Mach number, Reynolds number, and angle of attack. A review of the UPWT PSP and VMD systems is provided, and representative results obtained on selected HSR models are shown. A promising technique to streamline the wind tunnel testing process, Modern Experimental Design, is also discussed in conjunction with recently-completed wing deformation measurements at UPWT.

Erickson, G. E.↗
NASA NTRS1999

Pressure-Sensitive Paint and Video Model Deformation Systems at the NASA Langley Unitary Plan Wind Tunnel

Pressure-sensitive paint (PSP) and video model deformation (VMD) systems have been installed in the Unitary Plan Wind Tunnel at the NASA Langley Research Center to support the supersonic wind tunnel testing requirements of the High Speed Research (HSR) program. The PSP and VMD systems have been operational since early 1996 and provide the capabilities of measuring global surface static pressures and wing local twist angles and deflections (bending). These techniques have been successfully applied to several HSR wind tunnel models for wide ranges of the Mach number, Reynolds number, and angle of attack. A review of the UPWT PSP and VMD systems is provided, and representative results obtained on selected HSR models are shown. A promising technique to streamline the wind tunnel testing process, Modern Experimental Design, is also discussed in conjunction with recently-completed wing deformation measurements at UPWT.

Erickson, G. E.↗
NASA NTRS1979

Evaluation of improved algorithms for nonlinear transient structural heat transfer

The paper presents the evaluation of algorithms having potential for reducing computer time to integrate large systems of equations governing nonlinear transient heat transfer in complex structures. A series of test problems is being developed to test two algorithms. The Shuttle Orbiter frame model having 190 nodes was the subject of an analytical/experimental study. Plans are to perform calculations for the complete wing model having about 4000 nodes as well as a regional (plug) model having about 200 nodes.

Adelman, H. M.↗
NASA NTRS1940

Preliminary wind-tunnel investigation of an NACA 23012 airfoil with various arrangements of venetian-blind flaps

Report presents the results of an investigation made in the NACA 7 by 10-foot wind tunnel of a large-chord NACA 23012 airfoil with several arrangements of venetian-blind flaps to determine the aerodynamic section characteristics as affected by the over-all flap chord, the chords of the slats used to form the flap, the slat spacing, the number of slats and the position of the flap with respect to the wing. Complete section data are given in the form of graphs for all the combinations tested.

Wenzinger, Carl J↗
NASA NTRS1959

Low-Speed Wind-Tunnel Investigation of Blowing Boundary-Layer Control on Leading- and Trailing-Edge Flaps of a Large-Scale, Low-Aspect-Ratio, 45 Swept-wing Airplane Configuration

Blowing boundary-layer control was applied to the leading- and trailing-edge flaps of a 45 deg sweptback-wing complete model in a full-scale low-speed wind-tunnel study. The principal purpose of the study was to determine the effects of leading-edge flap deflection and boundary-layer control on maximum lift and longitudinal stability. Leading-edge flap deflection alone was sufficient to maintain static longitudinal stability without trailing-edge flaps. However, leading-edge flap blowing was required to maintain longitudinal stability by delaying leading-edge flow separation when trailing-edge flaps were deflected either with or without blowing. Partial-span leading-edge flaps deflected 60 deg with moderate blowing gave the major increase in maximum lift, although higher deflection and additional blowing gave some further increase. Inboard of 0.4 semispan leading-edge flap deflection could be reduced to 40 deg and/or blowing could be omitted with only small loss in maximum lift. Trailing-edge flap lift increments were increased by boundary-layer control for deflections greater than 45 deg. Maximum lift was not increased with deflected trailing-edge flaps with blowing.

Maki, Ralph L.↗
NASA NTRS1953

A correlation by means of transonic similarity rules of the experimentally determined characteristics of 18 cambered wings of rectangular plan form

The effects of one type of camber on the aerodynamic characteristics of rectangular wings at high subsonic and transonic speeds have been studied by applying the transonic similarity rules to the correlation of experimental data for a series of 18 cambered wings having NACA 63A2XX and 63A4XX sections, aspect ratios from 1 to 4, and thicknesses from 4 to 8 percent. The data were obtained by use of a transonic bump over a Mach number range of 0.6 to 1.1.

PROFILES - COMPLETE WINGS↗
NASA NTRS1951

A Comparison of the Experimental and Theoretical Loading over Triangular Wings in Sideslip at Supersonic Speeds

The results of an experimental investigation of the load distribution over two triangular wings in sideslip at Mach numbers from 1.20 to 1.79 are presented and compared with theory. The two wings tested have identical plan form, 45 degrees sweepback of the leading edge, and an aspect ratio of 4.0. One model was composed of round-nose airfoil sections and the other of sharp-nose, biconvex sections. For both wings the maximum thickness of streamwise sections was 6 percent and was located at the 30-percent chord.

DIHEDRAL - COMPLETE WING↗