Inaeronautics andaeronautical engineering,camber is the asymmetry between the two acting surfaces of anairfoil, with the top surface of a wing (or correspondingly the front surface of a propeller blade) commonly being more convex (positive camber). An airfoil that is not cambered is called asymmetric airfoil. The benefits of cambering were discovered and first utilized byGeorge Cayley in the early 19th century.^[1]

Camber is usually designed into anairfoil to raise its maximumlift coefficient C_Lmax. This minimizes thestalling speed of aircraft using the airfoil. An aircraft with wings using a cambered airfoil will have a lower stalling speed than an aircraft with a similarwing loading and wings using a symmetric airfoil.

One recent cambered design is called thesupercritical airfoil. It is used for near-supersonic flight and produces a higherlift-to-drag ratio at nearsupersonic flight than traditional airfoils. Supercritical airfoils employ a flattened upper surface, highly cambered (curved) aft section, and greater leading-edge radius as compared to traditional airfoil shapes. These changes delay the onset ofwave drag.

An airfoil is said to have a positive camber if its upper surface (or in the case of a driving turbine or propeller blade its forward surface) is the more convex. Camber is a complex property that can be more fully characterized by an airfoil'scamber line, the curveZ(x) that is halfway between the upper and lower surfaces, andthickness functionT(x), which describes the thickness of the airfoils at any given point. The upper and lower surfaces can be defined as follows:

${\displaystyle Z_{\text{upper}}(x)=Z(x)+\frac{1}{2}T(x)}$ 

${\displaystyle Z_{\text{lower}}(x)=Z(x)-\frac{1}{2}T(x)}$ 

An airfoil where the camber line curves back up near the trailing edge is called a reflexed camber airfoil. Such an airfoil is useful in certain situations, such as withtailless aircraft, because themoment about theaerodynamic center of the airfoil can be 0. A camber line for such an airfoil can be defined as follows (note that the lines over the variables indicates that they have beennondimensionalized by dividing through by the chord):

${\displaystyle \overline{Z}(x)=a\left[\left(b-1\right){\overline{x}}^3-b{\overline{x}}^2+\overline{x}\right]}$ 

An airfoil with a reflexed camber line is shown at right. The thickness distribution for aNACA 4-series airfoil was used, with a 12% thickness ratio. The equation for this thickness distribution is:

${\displaystyle \overline{T}(x)=\frac{t}{0.2}\left(0.2969\sqrt{\overline{x}}-0.1260\overline{x}-0.3516{\overline{x}}^2+0.2843{\overline{x}}^3-0.1015{\overline{x}}^4\right)}$ 

Wheret is the thickness ratio.

• Chord
• NACA airfoil
• Aerodynamic drag
• Zero-lift axis

Sources

• Desktop Aerodynamics Digital Textbook. Retrieved 9/7/08.
• Theory of Wing Sections, Ira H.Abbott and Albert E.Von Doenhoff (Dover Publications-1959)ISBN 0-486-60586-8