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A swept wing has a specified spanwise reference line inclined relative to the direction perpendicular to the aircraft's longitudinal axis. Aft sweep places the outboard part farther rearward; forward sweep places it farther forward. The leading-edge sweep angle and quarter-chord sweep angle can differ, so a sweep value needs a named reference line.^[ref]

A useful geometric construction shifts each spanwise strip of a wing planform fore or aft while retaining its local airfoil chord. If both edges receive the same shift, their separation and hence the integrated wing area remain unchanged. For an original constant-chord example with wingspan $b=12\mathrm{m}$ and $c=2\mathrm{m}$, define $x_{\mathrm{LE}}(y)=|y|\tan 30{}^{\circ }$. The tip leading edge is $6\tan 30{}^{\circ }\approx 3.4641\mathrm{m}$ aft of the root, while $S=bc=24{\mathrm{m}}^2$ and $AR=b^2/S=6$ are unchanged by the strip shift.

Preserving these geometric scalars does not preserve the three-dimensional flow. At high Mach number, the normal Mach number for a swept leading edge helps organize a simplified comparison of the component perpendicular to that edge. Finite-span flow, thickness and interference with the fuselage remain relevant.

The primary 1947 experiment by Richard T. Whitcomb used a different geometric change: it rotated the two wing semispans to obtain zero, forward and backward sweep, then obtained loading and drag information from pressure measurements and wake surveys. The shock-associated changes did not shift by exactly a universal cosine factor.^[ref]

Thus an experimental sweep comparison should identify what changed besides the angle: rotation of an existing wing, a stripwise shear of a planform, and a newly designed swept wing need not retain the same projected area or section arrangement.

Wing sweep describes the inclination of a chosen spanwise line in a wing planform relative to the direction perpendicular to the aircraft's longitudinal axis. It must name the reference line: a leading-edge sweep angle and a quarter-chord sweep angle can differ on the same aircraft wing. A swept wing therefore needs a reference convention, not just an unqualified angle.^[ref]

Let $x$ increase aft, let $y$ increase from root toward one tip, and define the straight leading edge and linearly varying airfoil chord byThe line at chord fraction $f$ has $x_f(y)=x_{\mathrm{LE}}(y)+fc(y)$. Differentiating gives the exact geometric relationFor $b=10\mathrm{m}$, $c_r=2\mathrm{m}$, $c_t=1\mathrm{m}$ and ${\Lambda }_{\mathrm{LE}}=20{}^{\circ }$, the quarter-chord sweep angle is $\arctan (\tan 20{}^{\circ }-0.05)\simeq 17.43{}^{\circ }$. This reference-line dependence of wing sweep is purely geometric.

There is also a useful high-speed approximation. Resolving the incoming velocity perpendicular to a swept straight leading edge gives the normal Mach number for a swept leading edge, $M_n=M\cos {\Lambda }_{\mathrm{LE}}$. Thus $M=0.9$ and ${\Lambda }_{\mathrm{LE}}=30{}^{\circ }$ give $M_n\simeq 0.779$. This component is a guide to compressibility effects, not an exact prediction of shock wave formation on a finite wing with thickness and a fuselage.

In a 1947 experiment, Richard T. Whitcomb compared unswept wings with forward and backward sweep in the Langley 8-foot high-speed wind tunnel. Pressure measurements and wake surveys over Mach numbers from 0.60 to 0.96 showed shock-associated changes occurring later for the swept configurations. The reported shifts were not exactly the simple cosine prediction.^[ref]