The static test involved inverting the aircraft with the nose angled 7° down and loading the wings with 1760lbs of sand bags which represents a 4.4g load. Thus the drag force = 1760(sin7) or 214.5 lbs. Is this correct or am I all wet?

The math is correct, but, what was angled at 7°? The airfoil nose or the airplanes nose? There is a couple or three deg difference there. So you could be looking at as much as 1760(sin10). Usually, high speed stall occurs at a Clmax that is close to (and maybe more than) the SECTION Clmax due to "dynamic overshoot", i.e., it takes an instant or two for the air to remember that it is supposed to let go.

1760/4.4 = 400lb, is this your gross wt?

As an example, the last (only, to be honest :whistle

aircraft that I calculated the chordwise forces on was an ultralight with Wmax of 500 lb and 119 ft² wing area. I fantasized that I could make this strong enough for aerobatics w/o busting the pt 103 weight limit. At 6g, the max chordwise force I calculated was 267 lb for ONE wing panel. The normal force was 1241 lb for one panel, so the "lift" vector would be √(1241² + 267²) = 1269 lb for one panel. 500 * 6 = 3000 lb. This is for the wing panel only and does not include the wing area passing through the fuselage and also includes inertial relief for the weight of the wing panel. So, on my design, the center section or cabane (if it had one, it was actually a mid-wing), would need to withstand 534 lb forward force. This is for a heavier plane at substantially more g loading than yours. These calcs were done for an alpha of 19.75 degrees. Dynamic overshoot and a conservative outlook are responsible for this high angle.

For structural design purposes, it is always good to be conservative, but not overly so. For example, from Hiscock's

*Design of Light Aircraft*:

"Chordwise loads, forwards, result from the chordwise component of the lift being greater than that of the wing drag component. At low angle of attack, any rearward component is too low to provide a design case.

However, drag loads should not be ignored. An arbitrary choice of a compareable rearward load will not produce a significant weight penalty."

Profile drag coefficients can be had from the airfoils drag polar chart, i.e., Cd plotted against Cl. The Minimax airfoil is a modified airfoil that Mr. Ison created and I gather that it has high drag at high alpha. Possibly similar to NACA 4415 or maybe even a little more.

Induced drag can be calculated by by Di = kL²/(pi*q*b²) where k=1.35 to 1.4 (it's 1.3 for a J-3 Cub) L is lift, q is mph²/391 (the standard Bernoulli equation rearranged) and b is the wing span in ft.

Add profile drag and induced drag, multiply that sum by cos(alpha), subtract from lift*sin(alpha) and this should give you your chordwise "forward" force.

Remember, the "static" test did not include drag force, so when all is said and done, your calculations might not be much different from the results of the static test, but I would encourage you to do these calcs anyway.