Editing Lift coefficient (section)

== Section lift coefficient ==
[[Image:Lift curve.svg|thumb|300px|right|A typical curve showing section lift coefficient versus angle of attack for a cambered airfoil]]
Lift coefficient may also be used as a characteristic of a particular shape (or cross-section) of an [[airfoil]].  In this application it is called the '''section lift coefficient''' <math>c_\text{l}</math>. It is common to show, for a particular airfoil section, the relationship between section lift coefficient and [[angle of attack]].<ref>Abbott, Ira H., and Von Doenhoff, Albert E.: ''Theory of Wing Sections''. Appendix IV</ref> It is also useful to show the relationship between section lift coefficient and [[drag coefficient]].

The section lift coefficient is based on two-dimensional flow over a wing of infinite span and non-varying cross-section so the lift is independent of spanwise effects and is defined in terms of <math>L^ \prime </math>, the lift force per unit span of the wing.  The definition becomes

:<math>c_\text{l} = \frac{L^ \prime}{q \, c},</math>

where <math>c\,</math> is the reference length that should always be specified: in aerodynamics and airfoil theory usually the airfoil [[chord (aircraft)|chord]] is chosen, while in marine dynamics and for struts usually the thickness <math>t\,</math> is chosen. Note this is directly analogous to the drag coefficient since the chord can be interpreted as the "area per unit span".

For a given angle of attack, ''c''<sub>l</sub> can be calculated approximately using the [[thin airfoil theory]],<ref>Clancy, L. J.: ''Aerodynamics''. Section 8.2</ref> calculated numerically or determined from wind tunnel tests on a finite-length test piece, with end-plates designed to ameliorate the three-dimensional effects. Plots of ''c''<sub>l</sub> versus angle of attack  show the same general shape for all [[airfoil]]s, but the particular numbers will vary. They show an almost linear increase in lift coefficient with increasing [[angle of attack]] with a gradient known as the lift slope.  For a thin airfoil of any shape the lift slope is 2π per radian, or π<sup>2</sup>/90 ≃ 0.11 per degree. At higher angles a maximum point is reached, after which the lift coefficient reduces. The angle at which maximum lift coefficient occurs is the [[Stall (flight)|stall]] angle of the airfoil, which is approximately 10 to 15 degrees on a typical airfoil.

The stall angle for a given profile is also increasing with increasing values of the Reynolds number, at higher speeds indeed the flow tends to stay attached to the profile for longer delaying the stall condition.<ref>{{Cite book|last=Katz|first=J.|title=Race Car Aerodynamics|publisher=Bentley Publishers|year=2004|isbn=0-8376-0142-8|location=Cambridge, MA|pages=93}}</ref><ref>{{Cite book|last=Katz|first=J|title=Low-Speed Aerodynamics: From Wing Theory to Panel Methods|last2=Plotkin|first2=A|publisher=Cambridge University Press|year=2001|pages=525}}</ref> For this reason sometimes [[wind tunnel]] testing performed at lower Reynolds numbers than the simulated real life condition can sometimes give conservative feedback overestimating the profiles stall.

Symmetric airfoils necessarily have plots of c<sub>l</sub> versus angle of attack symmetric about the ''c''<sub>l</sub> axis, but for any airfoil with positive [[camber (aerodynamics)|camber]], i.e. asymmetrical, convex from above, there is still a small but positive lift coefficient with angles of attack less than zero. That is, the angle at which ''c''<sub>l</sub> = 0 is negative. On such airfoils at zero angle of attack the pressures on the upper surface are lower than on the lower surface.