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Allison, D. O.

Publications and source records attributed to Allison, D. O..

14 records
All sourcesNASA NTRSDOE OSTI
NASA NTRS1987

EA-6B high-lift wing modifications

NASA-Langley has accomplished the computational design and experimental verification of EA-6B aircraft wing modifications for improved high lift capability. The modifications are comparatively simple, and attempt to improve low speed high lift performance while maintaining high speed cruise efficiency. Several two- and three-dimensional low speed and transonic computational techniques were employed, together with extensive wind tunnel tests. The modified inboard and outboard edge slat/flap system sections yielded efficiency improvements that were verified by three-dimensional wind tunnel experiments to amount to an 11-percent wing-body lift coefficient enhancement at low speed.

Waggoner, E. G.↗
NASA NTRS1985

Semiempirical method for predicting vortex-induced rolling moments

A method is described for the prediction of rolling moments on a wing penetrating a vortex velocity field generated by a large aircraft. Rolling moments are determined from lifting pressure coefficients computed with an inviscid-flow linear panel method. Two empirical corrections are included to account for the lifting efficiency of an airfoil section and the local stall on the wing. Predicted rolling moments are compared with those from two windtunnel experiments. Results indicate that experimental rolling moments, for which the Reynolds number of the following wing is low, should be interpreted with caution.

Allison, D. O.↗
NASA NTRS1983

Inviscid analysis of two supercritical laminar-flow-control airfoils at design and off-design conditions

Inviscid transonic flow results are provided at design and off design conditions for two supercritical laminar flow control airfoils. The newer airfoil, with its lower suction requirements for full chord laminar flow, has a higher design Mach number, steeper pressure gradients, a more positive pressure level in the forward region of the lower surface, and a recovery to a less positive pressure at the trailing edge. The two dimensional design Mach numbers for the two airfoils are 0.755 and 0.730 at a common design lift coefficient of 0.60, and their thickness to chord ratios are 0.131 and 0.135, respectively. Off design shock formation characteristics are similar for the two airfoils over a range of Mach numbers between 0.6 and 0.8 and lift coefficients from 0.4 to 0.7. The newer airfoil is similar to the one used in a large chord swept model experiment designed for the Langley 8 Foot Transonic Pressure Tunnel.

Allison, D. O.↗
NASA NTRS1980

Low-speed aerodynamic characteristics of a 14-percent-thick NASA phase 2 supercritical airfoil designed for a lift coefficient of 0.7

The low speed aerodynamic characteristics of a 14 percent thick supercritical airfoil are documented. The wind tunnel test was conducted in the Low Turbulence Pressure Tunnel. The effects of varying chord Reynolds number from 2,000,000 to 18,000,000 at a Mach number of 0.15 and the effects of varying Mach number from 0.10 to 0.32 at a Reynolds number of 6,000,000 are included.

Harris, C. D.↗
NASA NTRS1978

Design of a laminar-flow-control supercritical airfoil for a swept wing

An airfoil was analytically designed and analyzed for a combination of supercritical flow and laminar flow control (LFC) by boundary layer suction. A shockless inverse method was used to design an airfoil and an analysis method was used in lower surface redesign work. The laminar flow pressure distributions were computed without a boundary layer under the assumption that the laminar boundary layer would be kept thin by suction. Inviscid calculations showed that this 13.5 percent thick airfoil has shockless flows for conditions at and below the design normal Mach number of 0.73 and the design section lift coefficient of 0.60, and that the maximum local normal Mach number is 1.12 at the design point. The laminar boundary layer instabilities can be controlled with suction but the undercut leading edge of the airfoil provides a low velocity, constant pressure coefficients region which is conducive to laminar flow without suction. The airfoil was designed to be capable of lift recovery with no suction by the deflection of a small trailing edge flap.

Allison, D. O.↗
NASA NTRS1977

Comparison of experimental and theoretical drag characteristics for a 10-percent-thick supercritical airfoil using a new version of an analysis code

Comparisons of experimental and theoretical drag characteristics for a 10-percent-thick supercritical airfoil using a new version of an advanced analysis code. Comparisons are made at near-design normal-force coefficients for Reynolds numbers from 2 to 11 million. Comments are made concerning various input parameters to the code.

Harris, C. D.↗
NASA NTRS1976

Shockless airfoils with thicknesses of 20.6 and 20.7 percent chord analytically designed for a Mach number of 0.68 and a lift coefficient of 0.40

A 20.8 percent-thick airfoil shape was designed to have shockless inviscid flow at a Mach number of 0.68 and a lift coefficient of 0.40. In order to determine the actual airfoils which would yield this same shockless flow when viscous effects are included, boundary layer displacement thicknesses were subtracted from the inviscid shape for Reynolds numbers of 100 and 35 million. This process yielded airfoils with thicknesses of 20.7 and 20.6 percent, respectively. Subtraction of boundary layer displacement thicknesses for Reynolds numbers below 35 million yielded nonphysical airfoils, that is airfoils with negative thicknesses near tHe trailing edge. The pitching moment about the quarter-chord point at the design condition was -0.082 for the inviscid shape and, consequently, for both airfoils. Off-design calculations for the two airfoils were made using a computer program which provides for the interaction of the inviscid flow and boundary layer solutions. The pressure distributions of the airfoils were shockless for conditions from the design point to lower Mach numbers and lift coefficients. No boundary layer separation was predicted except in the last 3 percent chord on the upper surface.

Allison, D. O.↗
NASA NTRS1974

Comparison of interference-free numerical results with sample experimental data for the AEDC wall-interference model at transonic and subsonic flow conditions

Numerical results obtained from two computer programs recently developed with NASA support and now available for use by others are compared with some sample experimental data taken on a rectangular-wing configuration in the AEDC 16-Foot Transonic Tunnel at transonic and subsonic flow conditions. This data was used in an AEDC investigation as reference data to deduce the tunnel-wall interference effects for corresponding data taken in a smaller tunnel. The comparisons were originally intended to see how well a current state-of-the-art transonic flow calculation for a simple 3-D wing agreed with data which was felt by experimentalists to be relatively interference-free. As a result of the discrepancies between the experimental data and computational results at the quoted angle of attack, it was then deduced from an approximate stress analysis that the sting had deflected appreciably. Thus, the comparisons themselves are not so meaningful, since the calculations must be repeated at the proper angle of attack. Of more importance, however, is a demonstration of the utility of currently available computational tools in the analysis and correlation of transonic experimental data.

Newman, P. A.↗
NASA NTRS1972

Polynomial approximations of thermodynamic properties of arbitrary gas mixtures over wide pressure and density ranges

Computer programs for flow fields around planetary entry vehicles require real-gas equilibrium thermodynamic properties in a simple form which can be evaluated quickly. To fill this need, polynomial approximations were found for thermodynamic properties of air and model planetary atmospheres. A coefficient-averaging technique was used for curve fitting in lieu of the usual least-squares method. The polynomials consist of terms up to the ninth degree in each of two variables (essentially pressure and density) including all cross terms. Four of these polynomials can be joined to cover, for example, a range of about 1000 to 11000 K and 0.00001 to 1 atmosphere (1 atm = 1.0133 x 100,000 N/m sq) for a given thermodynamic property. Relative errors of less than 1 percent are found over most of the applicable range.

Allison, D. O.↗