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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 271 records · Page 15
All sourcesNASA NTRSDOE OSTI
NASA NTRS2009

Investigation of Seal-to-Floor Effects on Semi-Span Transonic Models

In an effort to achieve the maximum possible Reynolds number (Re) when conducting production testing for flight loads aerodynamic databases, it has been the preferred practice of The Boeing Company / Commercial Airplanes (BCA) -- Loads and Dynamics Group since the early 1990's to test large scale semi-span models in the 11- By 11-Foot Transonic Wind Tunnel (TWT) leg of the Unitary Plan Wind Tunnel (UPWT) at the NASA Ames Research Center (ARC). There are many problems related to testing large scale semi-span models of high aspect ratio flexible transport wings, such as; floor boundary layer effects, wing spanwise wall effects, solid blockage buoyancy effects, floor mechanical interference effects, airflow under the model effects, or tunnel flow gradient effects. For most of these issues, BCA has developed and implemented either standard testing methods or numerical correction schemes and these will not be discussed in this document. Other researchers have reported on semi-span transonic testing correction issues, however most of the reported research has been for low Mach testing. Some of the reports for low Mach testing address the difficult problem of preventing undesirable airflow under a semi-span model while ensuring unrestricted main balance functionality, however, for transonic models this issue has gone unresolved. BCA has been cognizant for sometime that there are marked differences in wing pressure distributions from semi-span transonic model testing than from full model or flight testing. It has been suspected that these differences are at least in part due to airflow under the model. Previous efforts by BCA to address this issue have proven to be ineffective or inconclusive and in one situation resulted in broken hardware. This paper reports on a Boeing-NASA collaborative investigation based on a series of small tests conducted between June 2006 and November 2007 in the 11 by 11 foot Transonic Wind Tunnel at NASA Ames on three large commercial jet transport configurations to assess the effects of sealing a semi-span model to the floor and to investigate efficient sealing and testing techniques. This document will show how sealing the model to the floor has a small but remarkably far reaching spanwise effect on wing pressures, wing local section forces and wing force summations.

Sleppy, Mark A.↗
NASA NTRS1956

A Factor Affecting Transonic Leading-edge Flow Separation

A change in flow pattern that was observed as the free-stream Mach number was increased in the vicinity of 0.8 was described in NACA Technical Note 1211 by Lindsey, Daley, and Humphreys. The flow on the upper surface behind the leading edge of an airfoil at an angle of attack changed abruptly from detached flow with an extensive region of separation to attached supersonic flow terminated by a shock wave. In the present paper, the consequences of shock-wave - boundary layer interaction are proposed as a factor that may be important in determining the conditions under which the change in flow pattern occurs. Some experimental evidence in support of the importance of this factor is presented.

FLOW, TURBULENT↗
NASA NTRS1978

Section drag coefficients from pressure probe transverses of a wing wake at low speeds

An in-flight wing wake section drag investigation was conducted using traversing pitot and static probes. The primary objective was to develop measurement techniques and improve the accuracy of in-flight wing profile drag measurements for low values of dynamic pressure and Reynolds number. Data were obtained on a sailplane for speeds from about 40 knots to 125 knots at chord Reynolds numbers between 1,000,000 and 3,000,000. Tests were conducted with zero flap deflection, deflected flaps, and various degrees of surface roughness, and for smooth and rough atmospheric conditions. Several techniques were used to increase data reliability and to minimize certain bias errors. A discussion of the effects of a total pressure probe in a pressure gradient, and the effects of discrete turbulence levels, on the data presented and other experimental results is also included.

Montoya, L. C.↗
NASA NTRS1983

Vertical drop test of a transport fuselage section located forward of the wing

A Boeing 707 fuselage section was drop tested at the NASA Langley Research Center to measure structural, seat, and occupant response to vertical crack loads. Post-test inspection showed that the section bottom collapsed inward approximately 2 ft. Preliminary data traces indicated maximum normal accelerations of 20 g on the fuselage bottom, 10 to 12 g on the cabin floor, and 6.5 to 8 g in the pelvises of the anthropomorphic dummies.

Williams, M. S.↗
NASA NTRS1978

A laser velocimeter flow survey above a stalled wing

A laser velocimeter operating in the backscatter mode was used to survey the flow about a stalled wing installed in the Langley V/STOL tunnel. Mean velocities and magnitudes of velocity fluctuations were calculated from measurements of two orthogonal components of velocity. Free shear mixing layers above and below a large separated flow region were defined. Velocity power spectra were calculated at two points in the flow field. The flow-field survey was carried out about a rectangular aspect-ratio-8 wing with an airfoil section. The wing angle of attack was 19.4 deg, the Mach number was 0.148, and the nominal Reynolds number was 1 x 1 million.

Young, W. H., Jr.↗
NASA NTRS1948

The effectiveness at high speeds of a 10-percent-chord plain trailing-edge flap on the NACA 65-210 airfoil section

This report contains the results of a high-speed wind-tunnel investigation of the effectiveness of a 10-percent-chord plain flap on the NACA 65-210 airfoil section. The results include an indication of the lift-producing characteristics and the effectiveness of the 10-percent-chord flap. From a comparison of the characteristics of the 10-percent-chord flap with those of a 20-percent-chord flap it was concluded that, although a reduction in flap-chord ratio from 0.20 to 0.10 lessens the severity of the effectiveness loss at supercritical speeds, the 20-percent-chord flap is more effective throughout the entire range of Mach numbers from 0.3 to 0.875.

CONTROLS, FLAP TYPE - WING SECTIONS↗
NASA NTRS1925

Calculation of wing spars of variable cross-section and linear load

The calculation of wing spars of constant cross-section and load has been thoroughly treated by a large number of authors. Such is not the case,however, regarding the calculation of wing spars whose section and linear load diminish toward the ends, as in wings of trapezoidal contour and decreasing section.

SPARS↗
NASA NTRS1991

Simulation of iced wing aerodynamics

The sectional and total aerodynamic load characteristics of moderate aspect ratio wings with and without simulated glaze leading edge ice were studied both computationally, using a three dimensional, compressible Navier-Stokes solver, and experimentally. The wing has an untwisted, untapered planform shape with NACA 0012 airfoil section. The wing has an unswept and swept configuration with aspect ratios of 4.06 and 5.0. Comparisons of computed surface pressures and sectional loads with experimental data for identical configurations are given. The abrupt decrease in stall angle of attack for the wing, as a result of the leading edge ice formation, was demonstrated numerically and experimentally.

Potapczuk, M. G.↗
NASA NTRS1991

Simulation of iced wing aerodynamics

The sectional and total aerodynamic load characteristics of moderate aspect ratio wings with and without simulated glaze leading edge ice were studied both computationally, using a three dimensional, compressible Navier-Stokes solver, and experimentally. The wing has an untwisted, untapered planform shape with NACA 0012 airfoil section. The wing has an unswept and swept configuration with aspect ratios of 4.06 and 5.0. Comparisons of computed surface pressures and sectional loads with experimental data for identical configurations are given. The abrupt decrease in stall angle of attack for the wing, as a result of the leading edge ice formation, was demonstrated numerically and experimentally.

Potapczuk, M. G.↗
NASA NTRS1992

Structural testing of the technology integration box beam

A full-scale section of a transport aircraft wing box was designed, analyzed, fabricated, and tested. The wing box section, which was called the technology integration box beam, contained blade stiffened covers and T-stiffened channel spars constructed using graphite/epoxy materials. Covers, spars, and the aluminum ribs were assembled using mechanical fasteners. The box beam was statically tested for several loading conditions to verify the stiffness and strength characteristics of the composite wing design. Failure of the box beam occurred at 125 percent of design limit load during the combined upbending and torsion ultimate design load test. It appears that the failure initiated at a stiffener runout location in the upper cover which resulted in rupture of the upper cover and portions of both spars.

Griffin, C. F.↗