I have an aircraft similar to a Cessna 172 and am using a virtual wind tunnel software (flowdesign) to get a drag estimate. I import the airplanes 3D model and I get around 2240 lbs of drag with a drag coefficient of 0.33 at a speed of 253fps, which I think is way too much to be trusted.
I then tried to calculate it using an equation (although the equation is supposedly only for wing drag): D=q∗S(Cd0+CL^2/(π∗AR))
Where q is dynamic pressure, S is the wing surface area (208sqft in my case) and AR is the wing aspect ratio (6). I used a D/q value of 7 (to be conservative) and a Cd0 of 0.033. With that, I got a result of 560lbs of drag (remember, this is supposed to be just for the wing) at a speed of 253fps (which is the speed at which the best L/D ratio is supposed to be according to the equation - and is the same speed I used in the virtual wind tunnel)
With that result, I then went to calculate how much thrust I would need for a climb rate of 1250fpm. The equation is: T=W*(Vv/V+D/L)
Where W is the weight of the airplane, Vv is the targeted vertical speed (1250fpm), V and D are the speed and drag for the best L/D ratio (so 560lbs and 253fps), L is probably the lift and I used the same value as the weight there. With that, I got the result of 923lbs of thrust needed. That is around 723 horsepower.
Now if I go look up the horsepower of a Cessna 172, google says it is 180. The climb rate is 720fpm, which is half of what I targeted for, but the difference is still too big for my calculations to be right. The wing area is 175sqft, only about 30sqft less than I have.
I then tried to calculate the drag with another equation, in which I additioned the induced and parasitic drag. The equation for the induced drag is: Di=Cdi*S*q
Where Cdi=Cl^2/(π∗AR) where Cl=W/(q*S) and I got a result of about 118lbs of induced drag. For parasitic drag, I used Dp=D/q*q and got around 284lbs of parasitic drag. The total drag is 403lbs, which I believe is still way too much to be true.
I then tried using the wright equation, which is: D=Cd*q*S, and I got a result of 3160 lbs of drag using a Cd of 0.033 (now I know I probably should have used something closer to what I got from the virtual wind tunnel, but the result would have been even greater) and a total area of 3160sqft (that is the total wetted area of the whole airplane).
All of the above results could be countered with enough horsepower (and that is somewhat realistically doable), but when I compare the horsepower required, the climb rate and the wing area to other already existing planes, my calculations seem off.
So what is a simple way of calculating the total drag of the entire airplane accurately?
Sorry for the wall of text.
P.S.: if anyone is interested, I am using these videos (or series of videos) to conceptualize an airplane design:,