Knockdown Factor For Cold Worked Material 1

[alexeu]

Consider this, a skin panel is sitting around waiting to be loaded on the assembly jig. However, it is found that approximately 1" x 5" of the edge has wrinkled due to mishandling of the skin panel. After cold working, and the wrinkling has been straighten, a conductivity test on the damaged area was performed. It is shown that the material conductivity at the damaged area is the same as the rest of the panel.

Question: Can the material properties and charatceristics of the skin panel at the damaged area be considered as the same as it was before the damage was inflicted on it?

 

[BATman2]

Alexeu,

Not an aircraft type and there are details I like to know but....

1) Assuming the skin hardenable material like 5000 series AL, I would say no, it has been locally hardened by the cold working. If its a heat treatable alloy, there should be no change.

2) No matter how flat you think it is, there is now an initial imprefection in skin, hence the buckling strength will likely be lower locally.

3) If this is a low stress area (if such things exist on aircraft), perhaps its not an issue. If its high stress, I'd get a new sheet.

My 2 cents,
Nick

 

[Sparweb]

Rarely is there enough quantitative information available to make a positive decision in a case like this. The situation is similar when there are dents in the skin, say from propeller ejecta.

Since you say that the damage is at the edge of the panel, then it is very likely that there are fasteners in this area attaching it to substructure, the adjacent skin panel, etcetera.

These fastener holes may not take well to being installed in a cold-worked area. For the same reason "edge distance" applies not only to sheet edges but to bends, installing the fasteners in an area with unspecified damage is also a bad idea. What manner of scratches and internal stresses were left behind by the original damage? Subsequent flattening may have restored the dimensions, but small zones that have been plastically deformed may buckle far too easily and undermine the strength of the overall panel. Worse, cracks that wouldn't have otherwise grown, now have a local high-stress area to draw them out of hiding.

It's easy to speculate, because there's no economical way to gather 100% information to make a perfect decision. On the other hand, the area is small, and let nobody fool you: this isn't the first time a sheet got wrinkled, flattened, and put on anyway. You have guts to admit it. Many would just shut up and pound it all together.


Steven Fahey, CET
"Simplicate, and add more lightness" - Bill Stout

 

[alexeu]

TO answer Batman 2, at times, it is not economical to get a new sheet due to time constraint unless numbers can be shown to substantiate the damage. As indicated by Sparweb, there are no conclusive method or 100% information to come to a perfect "engineering" decision.
In item (b) of Batman 2 response, to relieve the buckling load in the skin, alternate repair such as channels or hat sections can be installed concurrently on top of existing stringers that would be installed in the damaged skin vicinity. Thus relieving the load in the skin.
As indicated in my initial question, it is the edge of a skin, thus it is also the area to be used as lap splices. Theoretically, the fasteners should be picking up the load to be transferred to the splice member. Hence, buckling of the skin would not be my main concern, or should I be?? So the question would be, are there any documentation that shows fastener bearing in material that is imperfect or were there any test performed by any company to show bearing capability of fasteners in material that has known imperfections.
Sparweb, there would probably be gouges, nicks and scratches but those imperfection could be reduced by blending the damage area and putting 125Ra surface finish. Then have eddy current inspection to ensure no further damage. Of course, if the material removed is so high, then it warrant a replacement of skin. However, if it doesn't, how can engineers substantiate or determine the level of damage before replacing the damaged skin?

 

[Sparweb]

The engineer's typical answer applies here: "it depends".

You've gone beyong designing and building a part, and have entered into the realm of repairing it. The guiding philosophy is to "restore equivalent strength". Depending on the structure, the certification basis of the airplane, and the nature of the damage, you may also need to show "no degradation of service life".

The topic has been discussed to some degree, but not to the original poster's satisfaction in [thread16-6806] .

Where exactly does this skin panel go (wing/fuse/fairing), what is it made of (material/thickness/chem-milling), what rivets are involved (size/pitch/material), and the extent of the damage, (change in thickness, waviness, lost cladding) must all be taken into account before you START making your conclusion.

I will take exception to a small statement above that could lead in the wrong direction:

"to relieve the buckling load in the skin, alternate repair such as channels or hat sections can be installed concurrently on top of existing stringers that would be installed in the damaged skin vicinity. Thus relieving the load in the skin."

Just because you've added a stiffener, doesn't mean you've relieved a load. You can expect to stiffen the structure with it, but careful design is called for to prevent the stiffener from "gathering" loads from other structures, or transferring in more load than would otherwise be present.
One overloaded rivet at the end of your stiffener, in a thinned-out scratchy skin, could ruin somebody's day, years from now.

STF

 

[BATman2]

SparWeb,

Good response!! If the damage is truly localized, could a flat doubler plate be used without excessive stiffness? If the double spans over the joint, it would also address the joint issue.

I would be in favor of testing a mock up repair for static and fatgiue strength, assuming that's cheaper than getting a new panel.


Batman2

 

[alexeu]

Thanks Sparweb and BATman2 for your comments. I think the real argument here is; that if the imperfection (non-visible) is actually acceptable. I agree wholeheartedly with Sparweb's statement of restoring equivalent strength and no degradation of service life. The problem I posed initially was to figure out how to quantify the loss of strength and service life for such a damage.

As far as BATman2's second reply, I do not think a doubler would be a good idea because it is at a splice area. Thus, the build up of thickness at this part should not be encouraged due to eccentric loads which causes additional moments.A doubler would have to match the thickness of the skin or greater for equivalent strength. Thus, the thickness of the doubler could easily be greater than 50 thousand of an inch. If a stringer is located at the damaged area, you would also increase the stresses due to the preload by inserting a thick doubler at that location. For the sake of an argument I am being real general of the location of the damage, obviously a fairing would not be a big deal compared to the shell of the fuselage. Also, I am assuming Production environment which does not allow external doubler to be installed.

 

[Sparweb]

Since a production environment is being assumed, then I would ask for an ultrasonic inspection of the whole wrinkled/re-flattened area. The inspector should plot out a "map" of the remaining thickness, scratches, and pits that remain. Take the worst numbers, calculate the drop in inter-rivet buckling strength, panel buckling strength and - because I am also assuming you know the loads in the joint - the joint life given increased bearing stress at the rivets.

I would also steer clear of adding a doubler - that's not acceptable on a new airplane, nor is it likely to be necessary, once you've analyzed the impact of the damage.

STF