This is directed to anyone who may be interested and specifically to Mr. Teldorserious.
You know … throughout my participation on this forum, while it may have only been 8 years or so, I believe I have always tried to maintain some sense of professionalism with the others who participate here – even when they were off into some of their own beliefs – as limited as they were (and recognized as such by many other participants). So, if you would not object, I would like to suggest that you might consider reading the section of the rules that govern the certification of various kinds of airplanes.
In the US, those rules are written, published, and enforced by the FAA – in other countries those responsibilities fall to other organizations. As for the US, the section of the rules that govern the certification of transport category airplanes (like the A-300) is Part 25. This particular FAR Part addresses large civil airplanes and large civil helicopters. Transport category aircraft include jet-powered airplanes with 10 or more seats or with a maximum takeoff weight (MTOW) greater than 12,500 lb (5,670 kg); Propeller-driven airplanes with more than 19 seats or with a MTOW greater than 19,000 lb (8,618 kg); and Helicopters with a MTOW greater than 7,000 lb (3,175 kg). In the example you posted (an airliner and a “bizjet”) if the bizjet met the requirements to be certificated under the rules applicable to Transport Category jet powered airplanes, then, yes, both airplanes would be held to the same standard. You might wish to review the “yaw” standards posted in §25.351 - Yaw maneuver conditions. Here are those requirements:
The airplane must be designed for loads resulting from the yaw maneuver conditions specified in paragraphs (a) through (d) of this section at speeds from VMC to VD . Unbalanced aerodynamic moments about the center of gravity must be reacted in a rational or conservative manner considering the airplane inertia forces. In computing the tail loads the yawing velocity may be assumed to be zero.
(a) With the airplane in unaccelerated flight at zero yaw, it is assumed that the cockpit rudder control is suddenly displaced to achieve the resulting rudder deflection, as limited by:
(1) The control system on control surface stops; or
(2) A limit pilot force of 300 pounds from VMC to VA and 200 pounds from VC /MC to VD /MD , with a linear variation between VA and VC /MC .
(b) With the cockpit rudder control deflected so as always to maintain the maximum rudder deflection available within the limitations specified in paragraph (a) of this section, it is assumed that the airplane yaws to the overswing sideslip angle.
(c) With the airplane yawed to the static equilibrium sideslip angle, it is assumed that the cockpit rudder control is held so as to achieve the maximum rudder deflection available within the limitations specified in paragraph (a) of this section.
(d) With the airplane yawed to the static equilibrium sideslip angle of paragraph (c) of this section, it is assumed that the cockpit rudder control is suddenly returned to neutral.
I would call your attention to this last paragraph … where the airplane would be yawed to the static equilibrium sideslip angle and then the rudder control is suddenly returned to neutral. And we all should note that neutral is not full opposite rudder. While I am sure there are those on this forum who could provide chapter and verse and the formulas involved to calculate the inertia of the mass being “yawed” and what kind of force is generated on the vertical structure if full opposite yaw control (rudder pedal) were applied while the mass was moving in the original direction. If the airplane structure could withstand such reckless use of the rudder controls, all the better for the airplane – but to meet regulatory requirements, it is not required. There is no allowance for any transport category airplane anywhere (at least that I am aware of) that would not have a rudder application limit – and a cavalier “Van Damme kick to the rudder” at any time or place would likely result in that pilot’s immediate dismissal – assuming he and his airplane made it safely back to terra firma.
Tails of transport category airplanes do not “come off” at speeds under Va … or over Va … if the airplane is flown the way it was intended (and certificated) to be flown. In this particular case (AA587), the pilot flying successfully transitioned the first of two vortex encounters quite nicely – in fact, looking at the FDR, it was almost, if not, Textbook. The second vortex encounter was wholly different in that same pilot’s response. I have no idea if you have personally looked at any of the data applicable to this specific accident – if not, it would serve you well to do so … before you get to the level of conclusions that you’ve apparently reached without that kind of research/reading. However, and to alleviate your having to stomp around in knee-deep records, let me offer the following:
I’ll start by focusing on the initiation of the 2nd vortex encounter (and lest you wonder, my observations are directly from the FDR-generated display provided by the NTSB for this accident). The first control movement looks like it begins with a very small aileron correction to the left as the airplane looked to be ever so slightly beyond the 22-degree bank that seemed to be what the F/O was happy with during the departure turn as directed by ATC. Almost immediately we see a right rudder pedal deflection to the stop and a control wheel input of about 80 degrees to the right. The attitude indicator just prior to this looked to be relatively steady. Of course, with full right rudder and darn near full aileron to the right, the airplane begins a roll to the right (back toward wings level). However, when the airplane reaches about 20 degrees of left bank, while rolling to the right we see a simultaneous control wheel and rudder surface movement to the left. The rudder surface actually looks to exceed the pedal limit (but I don’t know if this is an anomaly with the indicator or not) and the control wheel gets to or awfully near to full control wheel displacement. This means at least 160 degrees of wheel change and probably something like 8 – 10 degrees of rudder surface change – all in 1 second, to the left as the airplane is rolling to the right. With all this opposite control surface input, the roll to the right is almost stopped (at about 10 degrees of left bank).
To add to the excitement, the full left aileron position is not maintained, nor is the full left rudder pedal position. The wheel is brought back to something like 10 degrees to the right simultaneously with application of full right rudder pedal deflection, again in 1 second, probably reaching constant rate saturation. While full right rudder is maintained, the control wheel is moved back to about 10 degrees left – again, in 1 second. As the rudder pedal deflection is maintained (very likely getting close to a stable sideslip), the control wheel is moved back to the right to just about full wheel travel and the rudder pedal exceeds the pedal limits (again, I don’t know if this is a display anomaly or if the actual limit was exceeded) – again taking a total of 1 second. Immediately, the rudder pedal is repositioned to full left deflection, and, in fact, goes well beyond the limits (again depending on the accuracy of the display), simultaneously the control wheel is deflected full left … again taking only 1 second. As the control wheel is moved back to the right (to about 45 degrees left), the rudder pedal deflection goes full right and the surface position presentation disappears, while the pedal position continues to show full right deflection. There is little doubt that this is where the data feed was stopped – probably because of the departure of the vertical stabilizer and rudder. The control wheel goes back to about neutral and back again to a right control wheel deflection of about 45 degrees.
The pitch attitude when this second event began was about 10 degrees nose up and the airspeed was 238 knots. By the time the rudder surface position display blanks out, there were 7 control wheel reversals and 5 rudder pedal reversals, all in about 7 seconds … and in this 7-second time frame the pitch goes to zero degrees and the airspeed increases to 251 knots.
Again, as I noted earlier, when an airplane is in a maximum equilibrium yaw, a sudden commanded full, or nearly full, opposite rudder movement against that sideslip can generate loads that exceed the “limit loads” and possibly the “ultimate loads” and can easily result in structural failure. I think the professional investigators reached this conclusion and if you have access to that NTSB produced video, you would likely see the same events and come to the same conclusion.