Warbird Information Exchange

What primarily causes fatigue cracks? Vibration. This is a preliminary investigation, so the FAA should be able to determine if there was another secondary factor other than vibration involved. To me, these would be total time of wing angle, and various types of corrosion. It'll be interesting to see what the final report says.

I'd guess exfoliation corrosion. The reason for this guess is based on past experience and the location. My friend ordered a stress panel from a guy in Arkansas, and guess what? Every single rib in it had exfoliation corrosion! Since this incident occured in Fla, it's likely this was the cause. If it was, there is already an AD on corrosion which applies to the attach angles too.

Vibration? (???) Try repeated stresses at or beyond limits,
and/or an improperly repaired section, or other loads or damage
presented to the part in question that (even if momentarily, over
time), exceeded design limits. Heat treated, aluminum airplane parts
don't last forever.... Corrosion can exacerbate the issue, but fatigue
cracks can form in aluminum parts that are pure as the driven
snow (no corrosion).

As far as drawing the conclusion that since the accident
occurred in Florida, it has to be because of exfoliation
corrosion... Eh? (???)

The aircraft in question had been registered in Illinois and
Ohio prior to moving to Florida (it spent the vast majority
of it's post-military career in Ohio). It didn't move to
Florida until 2000.

I'd like to know precisely what happened as much as the
next guy, but short of someone posting the metallurgical
report/findings here, we're collectively engaging in pure
speculation as to why the attach angle failed (due to
a fatigue crack as reported)....

Bela P. Havasreti

Which attach angle failed? The outer wing-lower attach
angle or the wing center seciton (lower) attach angle?

In the grand scheme of things, it probably makes no
difference, but curiosity gets the cat.

The parts in question are highly stressed, and need to
be looked after with care (not saying they weren't in
this case).

In my pile of parts, I have one wing center section that
has to be seen to be believed (the attach angles are
completely toast).

Does anyone know where the "wing-pull" AD came from
on the DC-3? When you look at a -3 wing, it's basically
a 1.5x scale model of the T-6 wing (constant chord
wing center section with out-board panels bolted on
with a bazillion bolts). This seemingly amazing coincidence
is solved by the fact that the same guy who designed the -3
moved from Douglas Aircraft to North American during
the time the T-6 came to be.

Bela P. Havasreti

Well, I'm not referring to this specific case, mine were general considerations, and if u understood so I'm sorry, and please dont put words in my mouth, I never meant to insult anybody here. No matter what I say u always think I'm offending somebody.. As per the "penguins" issue, i didnt follow the topic, but I think that the arrogant here is u, not the "penguins".. I'm sorry, i had the pleasure to exchange a couple of private messages with u in the Flypast forum, and I thought u were a different person, not the rude one I see here If u want to continue this discussion please send me a private message.
Now back to the original topic: do u think that a 200 hours inspection on the "L" junctions would be enough? Is there an average time for esfoliation and corrosion to occurr?

Cheers

Alex

I'm sorry if I offended you Alex, it's just that I have little patience for stupid, uninformed, or careless comments where warbirds are concerned.

If you haven't any first hand experience with this issue then it might be best to refrain from making, as you put it, "general" comments. These types of off the cuff statements are usually misinterpreted by those outside the warbird community and do nothing for those of us directly affected by this problem.

u prolly dont remember Steve, but I'm restorating a Harvard as well and this topic does interest me A LOT. I have seen a Cessna 170 and a Murphy Rebel fall here in Italy because of poor mantainance, that's why I made such adfirmations. I didnt mean to offend anybody really

Alex

... but I play one on TV.

Matt and Brandon, when I saw your comments, I had to go read the inspection process you referred to. I cannot believe it really says to remove bolts and then jack the outer panels. If they did not have cracks before the inspection, they are sure inviting them to occur by doing this. The attach angle was designed to spread the load over a wide area through the use of many bolts, each of which on its own does not withstand a very large load. The angle is neither very thick, nor very wide, but it works very well due to this design. Removing half the bolts will induce stress risers into each offending bolt hole, that it was never designed to take. Dye pening the attach angle while on jacks makes a lot of sense. Doing it with bolts removed makes no sense at all.

Another comment, and Matt, you hit the nail on the head. The T-6 has flown several million hours, with perhaps two documented inflight structural failures. This is an amazing safety record, especially when one considers that normal flight regime for a significant percentage of these flights has included aerobatics. To date, I have been very impressed and pleased with the FAA's methodical approach and restraint. Clearly, the airplane is very well designed and proven, but if it takes a dye penetrant inspection every 200 hours to ensure continued airworthiness, this T-6 owner/pilot will not complain.

_________________
Rob C

Historically, the claim of consensus has been the first refuge of scoundrels; it is a way to avoid debate by claiming that the matter is already settled. “

– Michael Crichton

Another thing to know would be about the airplane parking: was the plane parked in a open hangar? Is the area really humid or has relevant temperature jumps? The engineer who's following our restoration said that esfoliation is mainly found on planes left outside for years, where rain and all the weather agents might cause the esfoliation. He said that theorically esfoliation on a well preserved airplane is hard to find, unless there are some areas where where water condensation doesnt fade away.

Cheers

Alex

I wholeheartedly disagree that you can cause any damage jacking in this way.

http://www.boeing.com/companyoffices/doingbiz/environmental/TechNotes/TechNotes2000-02.pdf

Note the description of Stress Corrosion Cracking (SCC). Newer alloys (like 7050) are less prone to this than those from the 1940s.

"There are many ways to prevent SCC. Design loads on the aircraft
must not exceed material threshold stress levels for SCC."

"It occurs when there is a sustained tensile stress, and exposure to a corrosive environment."

After reading and posting earlier today, I got curious as to the official word on the subject of jacking with bolts removed. When I got home, I consulted my T-6 Handbook Erection and Maintenance Instructions (AN 01-60F-2). Mine was rev. 16 August 1949. Jacking is addressed on page 23 and various permutations are described. No where is there any reference to jacking with bolts removed.

Although we use military maintenance manuals out of necessity, one must remember that the T-6 is a normal category aircraft. Unless someone here can show otherwise, there is no approved data for the procedure. Any IA contemplating an inspection using this dubious procedure should ask themselves, in accordance with what am I doing this? I would stick to the manual until/unless the FAA releases an Airworthiness Directive which authorizes this procedure. One man's opinion, of course. The usual caveats.

_________________
Rob C

Historically, the claim of consensus has been the first refuge of scoundrels; it is a way to avoid debate by claiming that the matter is already settled. “

– Michael Crichton

There is a note (in the Erection & Maintenance manual I believe)
that says words to the effect of "Don't remove the fuel tank
cover / stress panels with the outer wing panels attached"
(for obvious reasons....).

I agree the design is stout as hell when everything is in place,
it's all in good shape, and all the bolts are in & torqued.

I'll repeat my question:

Does anyone have the specifics on the DC-3 "Wing Pull" AD?
The design is very similar (constant chord center section with
outer wing panels attached with a bazillion bolts through flanges).
I refer to my earlier post as to how this strange coincidence came
to be <grins>

I'd like to know if it was the center section (fuel tank stress
panel) lower angle that failed or the outer wing panel lower
angle.

Bela P. Havasreti

bdk wrote

I agree with brandon; Cracking is caused by cycling. Whether that cycling is from vibrations, or by cyclical stresses. It is caused by a frequency. Metal fatigues and cracks when it is moved back and forth over time.

bdk wrote

This true too, corrosion weakens metal, and if metal is weakened while onder a torsional, shear, tensile or compressive load; it will fail instantly.






That's conservative. However, I have no reason to believe that this will actually cause a crack. I wouldn't call it dubious, but I also don't know how it would help eitther. We don't need to rely on the FAA for everything. They set minimum standards only. If you want to go above their standard, you can go to an aerospace engineer, and have the parts in question analyzed using the latest stress analysis methods.

Bela wrote:



Yes, the T-6 wing is an exact scaled down DC-3 wing. From what I heard the guy at N.A. and Douglas was the one who designed both planes. This to me raises questions about the SBD, it has some remote similarities to the T-6. Such as wing attach angles in similar areas, similar landing gear, similar size, somewhat similar cockpit area. I'd bet that they were designed by the same guy, or another guy who was heavily influenced by the chief designer of the DC-3/T-6.

This question about the failure location has come up repeatedly. All of the information that has been public has clearly stated exactly where the evidence of failure was located.
refer to the aero-news article:
Preliminary examination of the wreckage shows a fatigue crack on the inboard lower center wing attach angle, which initiated in the radius of
the angle.
I have mentioned before on this board that it was the stress panel attach angle. I am now aware of two other T-6s that have been found to have corrosion by examining the attach points. These planes are now grounded pending replacement of the attach points. More information is coming out, but the investigation review has been moved to next week.

_________________
Live to fly, Fly to live.....

I had a long conversation with Fred from the FAA about the failure and was told that corrosion had nothing to do with the failure, and that it was a stress crack that took along time to devolop. The FAA seems to want a AD, even though it is not needed IMO. I am sorry it happened, but 2 failures is 65+ yrs is not a trend. He kept citing the South African inspection program and that they never had another failure, but he could not tell me how many they found to be cracked and the hrs in service of each found to be cracked. I believe the FAA is over reacting and the attach angles can be inspected as they all ready are, during a annual inspection. When I do a inspection, I look for corrosion, but I also go over the angles with a 10X glass. I have worked on a dozen or more various T-6 versions and have never found a attach angle cracked, I have found corrosion, which has caused me to replaced the angle.

Any added inspection on the attach angles, other than a visual inspection with a 10x or stronger glass is not needed, and will in the long run, cause more damage than it will detect. I have seen too many owners and A&P mechanics that did not know how to strip paint from aluminum correctly, I have seen them using steel wool, steel brushes, sanders, wiping paint stripper away and not netrulizing it with water. A large number of owners, if they have to strip the paint from the angles are going to leave them bare to make the next inspection easier, and that will open them for even more corrosion damage.

There is a more critical failure point in the T-6 which the FAA does not care about because no one has died, YET. The counter wieght bolt on the Ham standard prop has failed on 3 different planes I have worked on, and there was nothing in common, all had different times, were overhauled by different shops. 2 failed on T/O which almost caused the loss of the plane and crew, and 1 was caught before it failed.