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The Oswald efficiency, similar to the span efficiency, is a correction factor that represents the change in drag with lift of a three-dimensional wing or airplane, as compared with an ideal wing having the same aspect ratio and an elliptical lift distribution.
The ratio of the length (or span) of a rectangular-planform wing to its chord is known as the aspect ratio, an important indicator of the lift-induced drag the wing will create. [7] (For wings with planforms that are not rectangular, the aspect ratio is calculated as the square of the span divided by the wing planform area.)
In engineering, span is the distance between two adjacent structural supports (e.g., two piers) of a structural member (e.g., a beam). Span is measured in the horizontal direction either between the faces of the supports (clear span) or between the centers of the bearing surfaces (effective span): [1] A span can be closed by a solid beam or by ...
Note this is directly analogous to the drag coefficient since the chord can be interpreted as the "area per unit span". For a given angle of attack, c l can be calculated approximately using the thin airfoil theory , [ 6 ] calculated numerically or determined from wind tunnel tests on a finite-length test piece, with end-plates designed to ...
An ASH 31 glider with very high aspect ratio (AR=33.5) and lift-to-drag ratio (L/D=56). In aeronautics, the aspect ratio of a wing is the ratio of its span to its mean chord.It is equal to the square of the wingspan divided by the wing area.
a=chord, b=thickness, thickness-to-chord ratio = b/a The F-104 wing has a very low thickness-to-chord ratio of 3.36%. In aeronautics, the thickness-to-chord ratio, sometimes simply chord ratio or thickness ratio, compares the maximum vertical thickness of a wing to its chord.
Lifting line theory supposes wings that are long and thin with negligible fuselage, akin to a thin bar (the eponymous "lifting line") of span 2s driven through the fluid. . From the Kutta–Joukowski theorem, the lift L(y) on a 2-dimensional segment of the wing at distance y from the fuselage is proportional to the circulation Γ(y) about the bar a
The deflection at any point, , along the span of a center loaded simply supported beam can be calculated using: [1] = for The special case of elastic deflection at the midpoint C of a beam, loaded at its center, supported by two simple supports is then given by: [ 1 ] δ C = F L 3 48 E I {\displaystyle \delta _{C}={\frac {FL^{3}}{48EI}}} where