Observing how induced drag is derived, induced drag is loss of kinetic energy due to total vorticity that is produced by wing.
I want to find induced drag of finite wing with L=const over span.
Finite wing with constant lift distribution is just math concept,not existing in reality, cant be tested in any experiment because even wing has lots of twist,the wingtip area cannot satisfy Γ=const, too much pressure leaking.
This is formula for local induced angle of attack, if L=const over span, circulation Γ=const, dΓ/dy=0, implies alpha= 0 thus Dind.=0
-I am intersted will in Treffz plane get same result, has equation "dΓ/dy" inside?-
By definition airfoil has zero induced drag, I want to check this mathematically.
Airfoil is infinite wing with L=const over span, I think this integral (picture below) for alpha induced is valid. Same like in part 1 if dΓ/dy=0, Dind=0.
But when I use this formula: Di= L* 2 / (0.5 x ρ x π x V* 2 x b* 2) and put infinity instead b and L, (infinity wingspan will give infinity lift), I get Di=∞/∞
If I use L'Hospital's rule will I get Dind=0?
For specific case where L=W, indeed this formula get Dind=0, but I want to analyze in general case.