• the weight of the aircraft
• multiplied by the load factor (in this case, the 3.8Gs necessary to qualify the airplane for the "Standard" category)
• plus or minus the tail load for a particular flight condition (the center of gravity location was varied from most forward to most aft to determine the tail load to add [or subtract] from the wing load.)

That’s the total load. Now, how the load is distributed across the wing varies with the flight condition. Speed, angle of attack and control surface deflection all contribute to variations in distribution. It might be possible to put an unacceptable strain on one point of the wing without exceeding the total load. Loads are not constant from root to tip, nor from leading edge to trailing edge. From all the flight loads considered, we selected three "worst case" load conditions to test. If the wing is strong enough to survive these loads, then it will be strong enough in all the other flight conditions.

The first test case that we selected corresponds to the upper right corner of the V-n diagram in Fig. 1.

 

This is the result of a 3.8G symmetrical pull-up at 10 percent over redline airspeed. The wing angle of attack (AOA) for this case is 5 degrees. When the worst case tail download is considered, this puts the largest total load on the wing/center section and imposes the "worst-case" total bending load that the wing must carry. We design the wings to meet the standards of FAR Part 23 where this case is labeled "Condition D" so we have adopted the same terminology. You can get an idea of the spanwise load distribution for this case in Curve 1 of Fig. 2.



 Chordwise distribution is shown on Fig. 4.  Notice that the majority of the load on the forward one third of the chord.

The second test case that we selected corresponds to Condition A, (again, an FAA label we have adopted) the upper left corner of the envelope shown in Fig 1. This would result from a 3.8G symmetrical pull-up at maneuvering speed. While this is not the worst-case total load, it occurs at a 15 degree angle of attack. The spanwise distribution is shown on the upper curve of Fig. 2, but the interesting stuff happens elsewhere. At this AOA, the load on the wing forces it forward as well as lifting perpendicular to the chord plane. This results in a tension load on the joint between the rear spar and the fuselage. The chordwise load distribution for Condition A is shown in Fig. 3, where we can see that this condition results in most of the load being applied very near the front of the wing, which tries to twist the leading edge up. In engineerspeak it "places a large leading-edge-up torsional load on the wing." The wing is subjected to bending in two planes (forward and up) and twisting at the same time.

The third test case we selected corresponds to a symmetrical pull-up at two thirds of 3.8G at maneuvering speed plus full trailing-edge-down aileron deflection. You can’t read this case directly off the V-n diagram, but it is a good example of how combinations of loads must be considered. The wing angle of attack for this case is 10 degrees. The load for this condition is centered quite forward on the chord of the non-aileron portion of the wing (Fig. 3). The additional lift due to aileron deflection on the outboard portion of the wing (see the lower curve on Fig. 2) places a large bending load on the outboard wing. Because this lift is centered further aft (Fig. 5) it exerts a large trailing-edge-up twist as well.

We obviously can’t duplicate a dynamic situation in the shop, so we have to figure out the load distribution and simulate it with simple weight. If you superimpose the spanwise and chordwise load distributions on top of each other, you end up with a reasonably accurate picture of the total distribution of load over the entire wing root to tip and leading edge to trailing edge.

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