AOA and staying in the air

You can not believe how confusing this topic is and how much debate on flying : for now, i am going to put in a disertation from Dave Gall that i dont fully understand, but i think is correct.

The contribution of the fuselage is not likely to be significantly changed between the two fuselage attitude angles that you reported, nor are the moments generated by the changed heights of the drag forces significantly different between the two cases. Ignoring these factors (and any thrust effects) and investigating just the wing and canard, it should be easy to spot the contribution of the moments at the two different flight conditions. In fact, just general knowledge of normal cambered airfoils is sufficient to show that there should be a different fore-aft lift distribution just like the one you report from X-Plane:

Fact: most cambered airfoils have a negative pitching moment. Assume that it is a constant at all angles of attack (a not unreasonable assumption). In your first flight condition, you were showing 120 mph, and in your second you showed 180 mph. By simple ratios, it is apparent that the pitching moment of each lifting surface (wing and canard) is 225% greater at the higher speed, since pitching moment is proportional to the coefficient-of-pitching-moment (Cm) multiplied by the square of the speed. You increased the speed by 1.5, so you increased the pitching moment of each wing by (1.5)^2=2.25. This increased total-airplane pitching moment must be balanced by a change in distribution of lift between the two wings. But wait....

The effect is partially cancelled by the reduced pitching moment coefficient expected from the "reflexed" trailing edge and the reduced elevator deflection, but by how much? Certainly, those who argue that the lift distribution is ALWAYS EXACTLY the same because "the lift force" has to balance at the same location as the un-moved CG have confused the apples with the oranges (sorry, Ted). Their argument would have to consist of asserting that the reduced Cm of the reflexed trailing edges must necessarily always exactly cancel the increased pitching moment due to higher speed. That's a dubious connection at best, but also easily disproved: Every positively-pitch-stable airplane that flies requires an elevator displacement to fly at a lower or higher speed, thus discrediting the asserted always-constant lift distribution between the front and rear wings. Would these same people argue that the changed pitching moment of the TAIL exactly balances that of the wing at higher and lower speeds? Then what of all that we learned of tail-down-force in private pilot school? Are these people now asserting that the tail does its job by trying to TWIST the fuselage (in pitch) about the horizontal stabilizer spar? I think not.

But what's happening with the Raptor? At the higher speed, you show the rear wing being required to make MORE lift than had been required of it before. Apparently, the total airplane pitching moment is LESS at the higher speed, indicating that the effect on pitching moment of the reflex and reduced elevator deflection is MORE POWERFUL than the effect of the speed. Does this make sense? Yes. Whereas the speed would have increased the pitching moment by a factor of more than two, the retraction of a flap has the ability to reduce the pithing moment COEFFICIENT by a factor of five, ten, or more, or even to make it change sign (become positive). If you multiply a negative number by 2.25 then divide it by five, or ten, or even a negative number, the result is MORE POSITIVE than what you started with, so it makes sense that the total airplane pitching moment would be more positive at higher speed, thus loading the rear wing more and unloading the canard.

Does this relate to what is observed in conventional airplanes? Somewhat: retract the flaps on any spam-can and see if you don't also have to relax back-pressure to maintain altitude. You'll also notice an increase in airspeed. I think that's ample evidence of a positive-changed Cm. However, try flying that spam-can the way the Raptor is flown: dial in an elevator setting and clamp it in place (reflexor), then try to find the flap setting that keeps you out of the ocean. Dial in more trailing-edge-up on the elevator (reflexor) and you will quickly find yourself flying with the flaps in a different setting, but perhaps NOT more retracted as would be correct in analogy with tandem wing planes. This is because the flaps (elevators) on tandems are ahead of the CG whereas the flaps on conventional planes are aft of the CG.

That discrepancy notwithstanding, since Private Pilot school we have fooled ourselves into thinking that more up-elevator has meant more down force on the tail, but as we have just shown the opposite is actually the case. More up elevator means more down force INITIALLY, but once equilibrium is attained (either by flying slower or changing flap settings) the actual FORCE created by the horizontal stab/elevator on conventional planes is MORE POSITIVE ("less downforce") due to the changed airplane pitching moment. On tandems, the rear wing starts out lifting, so more positive means that it has to lift more. This condition is initiated by "more up elevator" on conventional planes, which translates to "trailing-edge-up reflexor" on tandems. Finally, equilibrium is attained on tandems by a changed "flap" setting (more trailing-edge-up on the canard). This gives the happy benefit of a FASTER speed due to reduced canard lift and drag.

To recap: Trailing-edge-up reflexor allows for faster flight by "unloading" the canard. This also puts more of the lift load on the rear wing, which suffers the disadvantage of having "retracted" its "flaps" to cause this condition. The rear wing thus flies at a higher angle of attack and has a lower critical (stall) angle of attack, so it flies with less margin below its stall angle. The canard flies with less load and has its "flap" (the elevator) more retracted, while flying at a slightly increased angle of attack, but not as much so as the rear wing (I won't bore you with downwash explanations). Thus, the canard has a larger margin below its stall angle of attack than it did previously, and is capable of larger multiples of its current lift load. That last translates into an increased load factor capability, or in other words, the up-reflexor gives the airplane a potentially LOWER maneuvering speed and stall speed. However, the rear wing is working harder and may be potentially able to stall BEFORE the canard stalls (is that your "wicked evil stall?").

I think it would be prudent to set VERY conservative limits on the allowable deflections of any reflexor, recognizing that at any given CG location, down reflexor can increase canard stall speed (relatively benign) and up reflexor may cause main wing stall (not nice). With that in mind, a careful envelope expansion would be needed to determine the useful/allowable reflexor positions for various CG locations, and the cruise/stall speed performance benefits of same.

In tandem wing designs, the maximum angle of attack (AOA) that the forward wing (canard) can achieve is greatly influenced by the deployment of its trailing edge. In most of the classical analysis literature, if you move anything on the trailing edge, it is called a flap. When you are talking about "canard" or "tandem wing" aircraft, the trailing edge of the front wing is called the elevator. Deployment of the "elevator" is typically given the Greek symbol of gamma.

The lift curve of the GU25-5(11)8 mod airfoil is greatly affected by the deployment of the elevator. For example, when the elevator is in the neutral trailing position (just hanging straight back) the max AOA of the airfoil occurs at a quite normal angle of 16 or 17 degrees. When the GU25-5(11)8 mod airfoil's elevator is deployed, however, the situation changes radically. At an elevator deployment angle of 20 degrees (gamma), the airfoil's max AOA is achieved at a very low 10 or 11 degrees. Any attempts to pull the nose up more only creates stall and pitches the nose down. This tendency severely limits the AOA of the fuselage (and the aft wing that is attached to it).

Click < here > to se a graphic of the effects of "gamma" vs "Stall AOA" for the GU25-5(11)8 mod airfoil. As the angle of deployment of the elevator (gamma) goes from 0 to 20, the stall AOA goes from 16 to 11 degrees.

The inevitable result of generating lift is to generate turbulence. At high AOA's, the forward wing of a tandem design generates a lot of turbulence in the form of shed vortex wake. This screwed up air trails backwards and will wash out any lifting surface behind it if it can. The trick, if you can do it, is to not let the aft lifting surfaces fall into the shed vortex wake.

As a tandem wing aircraft rotates about its center of lift, the aft wing tends to translate (vertical movement downwards) into the shed vortex wake sheet of the forward wing. If your forward wing is huge, like ours is, this is a very bad thing. Large, hard working wings generate a huge amount of disturbed air behind them. The fact that your aft wing is bolted to the fuselage only makes it worse. You, the pilot, are expecting the aft wing to keep on lifting its fair share of the load no matter what air it has to pass through. When the aft wing translates into the shed vortex sheet, its ability to create lift is greatly reduced. Lots of other bad stuff happens also, but the net result of pulling the nose up and screwing up the air for the aft wing is a very serious loss of aft wing lift. In its worst case, this loss of aft wing lift is called "deep stall". This is the condition were the forward wing is still lifting and the aft wing has completely stalled. This very depressing case allows the tail of the aircraft to sink down even more....pulling the aft wing even deeper into stall. From this condition, there are few recovery options.

Click < here > for a graphic interpretation of some "basic" data on the flight behavior of a typical Mark II D-Fly. Note : This is an interpretation of data given to us from a builder / pilot. The graphics show a GU25-5(11)8 canard with a 20% elevator as it was observed in flight. This is only a rough image of what is going on out there in the wind, so do not read more into the graphic than it shows. As the angle of attack of the forward wing is pulled to stall AOA, the shed vortex sheet over takes the aft wing.