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