Source: As noted earlier, C_{D,0} = C_D - C_{D,i}. The total drag coefficient can be estimated as: :C_D = \frac{550 \eta P}{\frac{1}{2} \rho_0 [\sigma S (1.47V)^3]}, where \eta is the
propulsive efficiency, P is engine power in
horsepower, \rho_0 sea-level air density in
slugs/cubic foot, \sigma is the atmospheric density ratio for an altitude other than sea level, S is the aircraft's wing area in square feet, and V is the aircraft's speed in miles per hour. Substituting 0.002378 for \rho_0, the equation is simplified to: :C_D = 1.456 \times 10^5 (\frac{\eta P}{\sigma S V^3}). The induced drag coefficient can be estimated as: :C_{D,i} = \frac{C_L^2}{\pi A\!\!\text{R} \epsilon}, where C_L is the
lift coefficient,
AR is the
aspect ratio, and \epsilon is the aircraft's
efficiency factor. Substituting for C_L gives: :C_{D,i}=\frac{4.822 \times 10^4}{A\!\!\text{R} \epsilon \sigma^2 V^4} (W/S)^2, where W/S is the
wing loading in lb/ft2. ==References==