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The normal Mach number for a swept leading edge is the incoming velocity component perpendicular to that edge divided by the local speed of sound. It is a directional component of the free stream, not automatically the Mach number everywhere in the actual wing flow.

Let $\mathbf{e}$ be a unit tangent along a straight leading edge, let $\mathbf{U}$ be the incoming velocity, and let $a$ be the speed of sound. Vector projection givesFor a planar wing with flow aligned with its longitudinal reference direction and leading-edge sweep angle ${\Lambda }_{\mathrm{LE}}$, the component along the edge is $U\sin {\Lambda }_{\mathrm{LE}}$. ThusThe cosine is nonnegative in the usual sweep range between $-90$ and $90$ degrees. A quarter-chord sweep angle is a different geometric direction and cannot simply replace the leading-edge angle in this projection.

For an original example, $M=1.2$ and ${\Lambda }_{\mathrm{LE}}=45{}^{\circ }$ give $M_n\approx 0.848528$. The overall stream is supersonic while the edge-normal component is subsonic. For $M>1$, the boundary $M_n=1$ occurs at $|{\Lambda }_{\mathrm{LE}}|=\arccos (1/M)$, about $33.5573{}^{\circ }$ for $M=1.2$. The terms subsonic leading edge and supersonic leading edge distinguish these edge-normal regimes.

Using the component as a predictor of an entire finite-wing flow adds assumptions. Thickness, finite span, angle of attack and interference with the fuselage can all matter. Whitcomb's pressure and wake measurements found that shock-associated changes with sweep did not follow the simple cosine shift exactly.^[ref] The projection itself is exact for the stated vectors; the reduction of a three-dimensional aerodynamic problem to that single component is the approximation.

A subsonic leading edge has an incoming edge-normal flow component below the speed of sound. It can occur even when the free stream is supersonic: the relevant component is the normal Mach number for a swept leading edge,Using a quarter-chord sweep angle instead would project onto a different direction and can change the classification incorrectly.^[ref]

For an original example, $M=1.6$ and ${\Lambda }_{\mathrm{LE}}=60{}^{\circ }$ give $M_n=0.8$. If the same aircraft wing is exposed to $M=2.2$ with unchanged sweep, $M_n=1.1$ and the edge is instead a supersonic leading edge. This is a flow-condition classification, not an immutable geometrical name.

For $M>1$ and aft sweep between zero and $90{}^{\circ }$, the threshold isAt $M=1.6$, the boundary is about $51.32{}^{\circ }$. Equivalently, the edge's angle from the downstream streamwise direction is less than the Mach angle, placing it inside the apex Mach cone. The normal Mach number for a swept leading edge article derives the underlying vector projection; no assumption about a measured surface pressure is needed for that geometry.

Subsonic edge-normal inflow does not imply that every part of the three-dimensional flow is subsonic or free of shock waves. Thickness, angle of attack and interference with a fuselage affect the surface flow. The regime is useful precisely when kept separate from those further aerodynamic predictions.