Regarding this (the original reason for the discussion):
Also... the laminar airfoil of the P-51 eliminates the increase of the coefficient of drag for the majority of the effective AOA. What that means is that as a pilot increases the AOA the coefficient of lift increases but the coefficient of drag remains constant for 65% (approximately) of the effective AOA. This is why a P-51 can raise its nose as much as eight degrees (or a little more) and the drag remains the same as if the wing were level and the P-51 can use its E to zoom to great advantage.
It dawned on me today why you might be thinking that. I wondered if you might be considering the section drag coefficient as being the same as the drag coefficient of the wing. Section drag coefficient is drag of the airfoil -- i.e., a wing of infinite aspect ratio. Once you have a wing of finite aspect ratio, it necessarily (even if the wing is in all other respects idealized) introduces induced drag that increases with lift (i.e, with angle of attack), and it is appreciable. There is no way around it -- it is a physical result of a wing of finite wingspan, whether the airfoil is laminar flow or not.
Let C_D = coefficient of drag.
For an idealized airfoil (infinite aspect ratio), C_D = c_d = constant, where c_d is the section drag coefficient (also "parasitic drag coefficient", "profile drag coefficient", etc. for the airfoil, and using various symbols, such as "C_D_0", "C_D_min", "c_d", etc. depending on reference). In the graphs I posted, it's called "section lift coefficient". In Pyro's graph, "parasite drag coefficient". In dtango's graph, "profile drag coefficient".
For an idealized wing producing constant value of downwash across the span (elliptical lift distribution is the optimal form here), C_D = c_d + C_D_i, where C_D_i = coefficient of induced drag = C_L^2 / (pi * AR), C_L is the coefficient of lift, pi is pi, and AR is the wing's apsect ratio. This is the way it is once one has a wing that is no longer infinite wingspan -- its a physical derivation that lift must induce drag of this form.
For wings that diverge from the optimal elliptical lift distribution, C_D_i = C_L^2 / (pi * AR) * (1 + delta), where delta is typcially small.
For real (non-idealized) airfoils, c_d does vary with C_L. For non-laminar-flow airfoils, c_d or section drag coefficient is roughly parabolic with C_L. For laminar-flow airfoils (ones where laminar flow happens for an appreciable amount of the chord), c_d is typcially better than for non-laminar-flow airfoils, and it can have a "drag bucket", where c_d is low and nearly constant for some amount of C_L before laminar flow is eventually lost and c_d goes back to being parabolic vs. C_L.
However, c_d is much less than induced drag beyond small angles of attack, so that isn't even needed to conclude major things about drag at 8 deg AoA. Thus, while I believe that for the P-51's airfoil c_d (or section drag coefficient) at 8 degress AoA is much greater than at zero deg. AoA (and you disagree), that doesn't even matter to the discussion of whether the P-51's wing has more drag at 8 deg. AoA than at zero AoA. c_d (i.e., C_D at zero lift) is on order 0.001 to 0.01 or so. C_D_i is much larger.
Let's look at some numbers. Let's (contrary to reality but for the sake of argument) take the P-51's airfoil to be 100% laminar flow and 100% the same as a perfect, idealized airfoil so that it's section drag coefficient is totally independent of C_L (or thus AoA) and equal to c_d_min. From the graphs I posted (and numerous other graphs of airfoil data), I'd guess that c_d_min for the P-51's airfoil is somewhere around 0.004, but also for the sake of argument, let's take it to be anything you want in the range of 0.001 to 0.01. Now let's look at C_D_i = C_L^2 / (pi * AR) (not even putting in any value for delta -- for the sake of argument letting the P-51 have an optimal elliptical lift distribution). The P-51's AR is about 5.8, so C_D_i = 0.0549 * C_L^2. At zero lift, C_D = c_d_min, which equals, say, 0.01. At 8 deg. AoA, where C_L is about 1.0, then C_D = 0.01 + 0.0549 = 0.0649 -- more than six times higher than C_D at zero lift. Or, if anyone wants to dispute that 8 deg AoA isn't a C_L of 1.0, fine -- put in something unrealistically low for the sake of argument -- say, 0.5. At C_L = 0.5, C_D = 0.01 + 0.0137, and C_D at 8 deg. AoA is still more than two times what it is at zero lift. In all cases, I would call that "substantially more".
References for the above, so that (although it is standard material in aerodynamics books), I can't be accused of making it up or misderiving it.
Airplane Performance, Stability, and Control, by Perkins and Hage
Aerodynamics, Aeronautics, and Flight Mechanics, by McCormick
Fundamentals of Flight, by Shevell
Theory of Flight, by Von Mises
Introduction to Flight, by Anderson
