HN39 asked why I thought the material of part "arm 36g" mattered:
The NTSB report says about the equivalent part of the A300: "The support strut and its attachment fitting are made of an aluminum alloy."
In the context of Bearfoil's structural preferences, I would like to understand why the material is important. Whatever its material or method of manufacture, that part needs to be sized to support an ultimate load of 120,000 N, and a limit load of 80,000 N. The important evidence is that those design loads were exceeded.
Thanks for the information that it is an aluminum bar. I read the 587 report (after 447 Report #1) but either missed that, or did not attach enough significance to it at the time to remember the detail.
Metal structures are a bit better understood, having a longer history. That said, their design requirements tend to be writ in rather round numbers. For example, I could correctly say that a suddenly applied impact load would cause a stress in this bar twice that of the stress resulting from the same load gradually applied. That is, an 18 g load, applied as an impact, would fracture this bar.
But then there are 2 bars, and one probably bears on its bolt (these must be individual capscrews because of the divergent angle between the bars) before the other. There is in some applications a factor to reduce the effectiveness of bolt groups because of this; IIRC it is applied as a reduction in tabulated permissible bolt loading. I forget if it is a reduction to 90% or to 80%; let's say 80% in this example, to keep the math simple-- I want to illustrate the nature of the issue with somewhat real numbers, but not split hairs over any exact number. So now, a 15 g impact load would fracture one of the two bars.
Now the hole in the bar is a stress raiser. That factor can easily be two or three, although I don't think it is necessarily as high as 3-- we just can't see under the bolt head to judge (it may be an oversized washer-faced head) the hole size. This stress appears as a linear singularity rising to that value along the inside surface of the hole bore. Say the factor is 2. Then a 7.5 g impact load would raise the stress along this line to ultimate strength of the material, and start a crack.
But a fatigue crack might start at the limit load, given enough repetition. That is, fatigue cracking is a factor to inspect for if 5 g impact loads can result from rudder operation. I don't know at the moment; it is just a question to ask. There is a crack formed on each end of a diameter at right angles to the axis of the bar (but see below). If these cracks progress 1/3 the way into the aluminum on either side of the hole, a 5 g impact will break one of the bars. If they have only gotten half that far, a 10 g uncharacteristic (let's say) impact load will break one of the bars.
So you see that the 24 g limit load required by the US regulation is a sort of semantic thing. It is arrived at by multiplying some known impact loading determined in the distant (I assume distant) past to have caused trouble which was judged economically avoidable. It's because a certification test would not apply an impact load; the test setup would apply load at a steady rate over a specified number of minutes (say 2 to 5, as the technician has to take notes-- or did in my time).
So the limit load required is not an impact load; the part is new, so fatigue cracks are not present; metallic ductility may reduce the effect of stress concentration at the bolt hole to some extent; the requirement for one bar may be simply set at half the requirement for two-- remember regulations are a legal document, not an engineering lesson book.
Similarly the ultimate load factor, the additional "capacity" amounting to half the limit load, is intended for other intangibles-- things difficult to quantify, either at all, or at reasonable design or fabrication cost. These could be random defects in the metal, burring of finish on surfaces subject to fatigue, bolts not filling the bolt holes, bolts not concentric with the bores; uneven stress distribution over the bar section below the taper reduction in thickness; secondary (bending) stressing caused by corrosion binding the bolts in the bolt holes; and so forth. If you figure it, it is in the 24 g's; if not, it's in the 12 g's above that-- just a rule of thumb.
Not a very precise science, is it? And the more complex the object to be tested, the more costly the test; the fewer tests likely to be run; and the less likely the results will be repeatable from one test to the next. Well, the VS itself is an example of such a complex object.
This is just a general explanation. I'm sure that HN39 understands all this. I'd just say that I could not have made such a clear explanation had the link arm been fabricated from a composite material. I'd explain further, but it's a big job which I should not tackle late at night. Anyway, that's why I wanted to find out what the material was.
BTW, I did finish a further explanation late last night, and realized in doing it that I knew why the aluminum bar "36 g" had broken. It broke in lifting the rudder as the VS broke loose to somesault forward over the fuselage. The sequence is (a) VS breaks loose in carbon-composite over rear anchor set; (b) VS breaks loose at middle anchor set, pulling it out due to the brief greater strength of the carbon-composite under impact loading; (c) this frees the VS to rotate forward, accelerating; (d) the aluminum bar "36 g" breaks as it attemps to keep the rudder pintles properly aligned while the pintles lift the rudder in this sudden rotation that also starts as an impact. Far less than 36 g's of upward acceleration is required to break this bar. I have the detailed explanation written out but it is too long for this post (and the margin isn't big enough to hold it either). I'll post it latter.
OE