= 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.<Wind-tunnel Investigation of Effects of Rear Upper Surface Modification on an NASA Supercritical Airfoil by Charles D. Harris and James A. Blackwell, Jr.>[Introduction, printed pp. 1–2]{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.<Wind-tunnel Investigation of Effects of Rear Upper Surface Modification on an NASA Supercritical Airfoil by Charles D. Harris and James A. Blackwell, Jr.>[Summary; Apparatus and Techniques, printed pp. 4–6]{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 $U=230\,\mathrm{m/s}$, 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.
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