Permanent or temporary fix
 

Would anyone hazard a guess if this would be a permanent repair or a temporary fix to get the aircraft to one of the two assembly plants?

If I would have been involved in the repair? At any time!
I did already aerobatics on a composite glider where the wing was in 3 pieces before we repaired it. A scarf joint repair of unidirectional carbon fibre spar caps works fine. Why should a wound fibre fuselage be different ?

Let´s hope the gale forcasted for tonight will not be stronger than the VS stand... Otherwise we have the next composite repair :ouch:

VFD
 

The tent is to keep the workers dry and walm and dirt out of the repair. It is NOT part of the temp critical post cure prosess.

What is so pleasing about the posts above is the total lack of understanding shown by most above about composite repair techniques and the scale of what can be done, on the basis of this I feel that my income over the next few years is very secure.

Good article here in the Seattle Times about the repair technique:

Boeing readies patch for fire-damaged 787 | Business & Technology | The Seattle Times

Apparently it's scheduled to take 5 weeks, and will be followed by flight tests of the aircraft while fitted with strain gauges before the repair can finally be signed off.

ET expect to have the aircraft back in service in about two months' time.

...not to mention the tent!

DWS:
Excellent point.:ok:

A source of valuable new data, which hasn't cost any lives. Not forgetting, the ELT lesson for Honeywell.

Deleted wrong thread

Hopefully not!
It would be much more reasonable to cut skin damaged by the fire probably in a rectangular cut and scarfen the edges, then to cut out the patch several inches larger than the hole, scarfen the edges and bond it in place with an epoxy resin which will transfer the loads as the fuselage is pressurized and unpressurized...

because otherwise indeed

Pardon my (already demonstrated) ignorance of composite repair techniques, but what does to "scarfen the edges" mean ?

A scarf joint is often used in wood work, e.g. for hat boxes and drums. Each end is "scarfed" with a diagonal /overlapping face, so when bent into a circle there is consistent thickness and strength, but a long diagonal overlap so that there is a large surface area for glue to hold to.

Volume's idea seems totally sound to me, but whether it is normal or needed with composites is beyond my knowledge. If the taper of the scarf joints
were in the orientation in which the panel would "plug" into a "socket" on the inside of the fuselage it would thus be almost impossible for it to blow out even in the event of serious adhesive failure.

On a lighter note, the breeze is steadily increasing as predicted, so I am hoping the Queen of Sheba's tail does not blow off with the wind.
(Apologies to Peter Cook and Dudley Moore).

It’s when you in essence chamfer the mating faces of a butt joint to increase the surface (bonding) area to achieve a higher strength repair. :ok:

how thick is the 'panel' here?

I assume we are talking <2mm? so not much point in "scarfen the edges"?

Actually it is not primarily to increase the bonding area (compared to a simple overlap), but to adopt the parts thickness to the loading, so that the deformation remains constant over the length of the bonding joint, hence the shear stress in the bond is almost constant. In a simple lap joint you would have enormous stress peaks at the ends of the bond joint, and the center would not carry any loads.
You smoothly transfer loads from one item to the other and you proportionally transition the wallthickness from full at the begining to zero at the end of the joint.

I noticed that the report said that the edges of the panel and the hole in the fuselage would be rounded. Would they both be convex with an "hour glass" section of sealant, concave with a "circle" section of sealant, or a convex meeting a concave and a unison thickness of selant?

I believe I did spot the word "sealant" and assume this would be a specialist adhesive/resin rather than bathroom sealant!

Joyride
 

I think the report you have seen is wide of the mark by some considerable distance.

I have nothing to add to the very good posts by Volume.

The Boeing method seems to be very similar to what has been done with GRP boats for decades, although in a different order, for relatively small damage to hulls.

With a boat, you clean up the hole/damage, usually squaring it off, and then bond a patch on the inside of the hull, but at least 4X its area , so that it overlaps the hole by a lot. The patch is the same thickness as the hull. The patch must follow the curve of the hull precisely, very difficult with a double curve. This patch is the strength of the repair.

A second patch, also the same thickness and curve as the hull but fitting the hole very accurately, is then bonded into the the hole itself. A really skilled worker will reverse scarf it if possible.

The gap is then made invisible with gelcoat matched to the hull.

Insead of pre-manufactured patches, a yard may plug the hole and then lay several layers of GRP to the same strength as the hull over the plug and surrounding GRP, then remove the plug and when the first stage has cured lay more GRP on the outside, in the hole, finishing off with gelcoat over the whole repair. This method is best when the curve makes a pre-made patch impractical.

With a boat, the pressure is from the outside, below the waterline. With the aircraft it is from the inside so this method would be even stronger.

No first hand knowledge of the area in question, but based on the composite fuselage pieces I have, seen I'd say 1/4 inch minimum (~6 mm), probably closer to 3/8 inch (~9 mm). Granted, I never actually measured anything, but visually I was a bit surprised at how thick they were (then again, I had a similar reaction the first time I saw an aluminum fuselage skin - they are thicker than most people think). Remember, unlike that access panel that fell off in India, we're talking primary structure here.

Capot
 

Your boat repair may work and deal with the load transfer but at massive weight increase, as for filling in the gaps with gel coat I think that is just setting up for further problems in the future.

The whole point about the repair techniques so well discribed above by Volume is that the repair when finished is exactly the same as the original structure both in load transfer and weight.

What we are seeing above with all this talk of patches is a metalcentric view to load transfer and an assumption that the " glue" joint is weaker than the original structure, this is simply untrue, the "glue" or more correctly resin IS exactly the same as the original structure and so correctly installed will react exactly like the original structure. Therefore no additional patches or plates are required as long as a sufficient scarff area is avalable.

Another airline has already had a skin puncture repaired as described, clean up the damage to a defined shape (in this case it was an L shape) scarfed edges, drop in a plug of the same dimensions.
You can't see the join. Admittedly, this Ethiopian job is much bigger but the principle seems the same.

A & C
I intended only to remark on the similarity of the approach developed over the decades for GRP boats and the method devised by Boeing for its aircraft.

There is NO suggestion that the standard required is remotely similar in any respect, or that the details of the process are or even could be the same!