= Subsonic leading edge
{title2=$M_n<1$}
A \b[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>,
$$
M_n=M|\cos\Lambda_{\rm LE}|<1.
$$
Using a <quarter-chord sweep angle> instead would project onto a different direction and can change the classification incorrectly.<Introduction to Aircraft Performance and Static Stability by Earll Murman>[PDF p. 9, Basic Sweep Concept, normal Mach component and subsonic leading-edge regime]{ref}
For an original example, $M=1.6$ and $\Lambda_{\rm LE}=60\,{}^\circ$ 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 is
$$
\Lambda_{\rm LE}>\arccos(1/M).
$$
At $M=1.6$, 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.
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