 Supercritical airfoil

A supercritical airfoil is shaped to control transonic flow so that shock-related drag rises less severely over its intended operating range. Here supercritical refers to operation beyond the onset of locally sonic flow around the section; the freestream need not be supersonic. Characteristic shaping includes reduced upper-surface curvature over much of the chord and substantial aft camber, rather than a requirement that no shock ever appear.

A primary NASA experiment describes the design objective as reducing losses from shock waves and flow separation, using a large leading-edge radius, a flattened upper surface and a strongly cambered trailing region. It investigated changing the rear contour to suppress an off-design second velocity peak.^[ref] The pressure recovery must work with the viscous boundary layer; a favorable change at one condition can worsen another.

The apparatus makes that tradeoff observable. Tests in the Langley 8-foot transonic pressure tunnel covered freestream Mach numbers $0.60$–$0.81$. Flush surface orifices supplied pressure distributions, while a rake about one chord downstream sampled wake pressures for drag inference. Controlled transition strips helped specify the boundary-layer state. The tested attempt to reduce the second peak adversely affected drag at design conditions.^[ref] These observations motivate off-design pressure recovery of a supercritical airfoil; optimizing a single contour feature is not enough.

For an illuminating coefficient calculation, define $C_p=(p-p_{\infty })/q_{\infty }$. A measured pressure recovery from $C_p=-0.8$ to $-0.3$ over a short distance has $\Delta p=0.5q_{\infty }$. At stipulated $q_{\infty }=12\mathrm{kPa}$ this is $6\mathrm{kPa}$. The abrupt recovery identifies a region to investigate with higher spatial resolution and a compressible-flow model; those two numbers alone do not establish the full shock strength or separated-flow extent.

A separate original performance comparison fixes reference area, dynamic pressure and lift coefficient. Reducing measured drag coefficient from $0.030$ to $0.026$ at $q_{\infty }=12\mathrm{kPa}$ and $S=20{\mathrm{m}}^2$ reduces drag by $960\mathrm{N}$. At $U=230m/s$, the associated drag-power difference is $220.8\mathrm{kW}$. These are stipulated values illustrating how a coefficient improvement translates into force and power; they are not a promised improvement for every supercritical section.

A useful comparison therefore matches Mach number, Reynolds number, lift, surface condition and transition treatment, and measures both surface pressure and wake or balance drag. Compare the full intended operating range, including lower-speed lift and off-design conditions. Thin-airfoil incompressible calculations cannot by themselves establish transonic drag performance.