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Thinking - Very Dangerous!

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Topaz

Super Moderator
Staff member
Joined
Jul 29, 2005
Messages
15,993
Location
Orange County, California
I've decided to go back and review my various spreadsheets for my project and I've stumbled across something that's confusing me.

It may be just the late hour :zzz: , but...

For my various performance calculations, total aircraft drag is obviously an important number. I've been generally using Raymer's methods, from his textbook.

He's defining the drag polar in the conventional fashion:

Cd(total) = Cd0 + Cdi*CL^2

where Cd0 is the parasite drag of the entire airframe and Cdi is the induced drag coefficient.

Then he moves on to pulling total aircraft drag from this, which is where I get confused:

Drag(total) = Cd(total)*q*Sref

where Sref is the reference wing area.

Why are we defining total drag in terms of the wing area only? Shouldn't the total drag equation look something like this?:

Drag(total) = Cd0*q*Swet+Cdi*CL^2*q*Sref

that is, add the total parasite drag (for the whole wetted area) to the induced drag of the wing alone?

I checked both his text and the example problem (the aerobat) and he's definitely doing it the former way. What am I missing?
 
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