**Related Resources: fluid flow**

### Aerodynamics Airfoil Theory Equations

**Aerodynamics Airfoil Theory Equations **

The lift force on an airfoil is given by

C_{L} = the lift coefficient

V = velocity (m/s) of the undisturbed fluid and

A_{p} = the projected area of the airfoil as seen from above
(plan area). This same area is used in defining the drag
coefficient for an airfoil.

The lift coefficient can be approximated by the equation:

C_{L} = 2 π k_{1} sin ( α + β ) which is valid for small values of α and β

k_{1} = a constant of proportionality

α = angle of attack (angle between chord of airfoil and
direction of flow)

β = negative of angle of attack for zero lift.

The drag coefficient may be approximated by:

C_{D∞} = infinite span drag coefficient

The aerodynamic moment is given by:

where the moment is taken about the front quarter point of the airfoil.

M_{C} = moment coefficient

A_{p} = plan area

c = chord length

Reference:

Fundamentals of Engineering, seventh Edition National Council of Examiners for Engineering amd Surveying

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