Simple Aerodynamic Equations

Lift Coefficient

CL = 2L / (⍴ v² A)

Drag Coefficient 

CD = 2D / (⍴ v² A)

Drag Coefficient for an Aircraft

CD = CD0 + CDI = 2D / (⍴ v² A) + CL² / (π AR e)

Aerodynamic Efficiency

L/D = CL/CD


CL = Coefficient of Lift
CD = Coefficient of Drag
L = Lift
D = Drag 
⍴ = Density
v = Velocity
A = Wing Area
AR = Aspect Ratio