Aerodynamics Airfoil Theory Equations

Related Resources: fluid flow

Aerodynamics Airfoil Theory Equations

Aerodynamics Airfoil Theory Equations

Open: Airfoil Lift and Characteristics Calculator

The lift force on an airfoil is given by

lift force

C_L = the lift coefficient
ρ = fluid density
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 is accurately determined by test experimentally. The lift coefficient can be approximated by the equation:

C_L = 2 π k₁ sin ( α + β ) which is valid for small values of α and β

alternatvely

C_L = 2 L / ( A ρ V² )

k₁ = 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:

The drag coefficient

C_D∞ = infinite span drag coefficient

infinite span drag coefficient

The aerodynamic moment is given by:

aerodynamic moment

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

M_C = moment coefficient
A_p = plan area
c = chord length

Airfoil

Reference:

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