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!

Typically we scarfen unidirectinal CFRP 1:50, so that would give a 100 mm scarfened overlap.
Depending the exact layup of the barrel the value might be a bit lower, as it for sure is not fully unidirectional.
Some Information from a very old glider repair handbook (all glassfibre)

A and C: Why shouldn't it bother me that there's no continuous carbon fiber running through/across the new joint?

Poorjohn
 

In answer to your question.............Because a properly executed scarf joint replicates the load transfer property's of the original structure.

When doing such repairs normally a test sample is made and sent to the aircraft manufacturer for destruction testing, we have never had a test sample fail to meet the new specification and on average the test samples are 1-3 % stronger than new items.

One way to think of this is to consider 2 hypothetical joints:

At one extreme would be a "butt" joint with resin (glue) in the gap, in this case there would be no fibers "across" the joint.
BTW: Does not matter if the resin or composite is stronger (actually stiffer) this creates a discontinuity / stress concentration aka not a good thing.

At the other extreme is a joint made by taking two barrel sections and tapering each from full thickness to none over the full length so one exactly fits inside of the other.

In this case there will be continouse fiber over the entire joint, although none transitions from one section to the other.

A scarf joint is similar to the second case expcept for the overlap distance.

BTW: Although I have some understanding of stress and transitions etc I am not a composites expert (at all) so feel free to correct/clarify the above. I am writing this partly to clarify my own understanding.

Scarf joints in wood (same situation, no continuous fibre, but substantial overlapping of parallel fibres) have been successfully used in both new construction and field repair. For perhaps a century!


BTW, the Howard DGA-8 has a Vne of 250 kt.

Well actually even longer than that! The burial ship at Sutton Hoo which is thought to date from around 620 AD was found to use splayed scarf joints. I think that is thought to be the oldest known use in the UK but that means they were probably used elsewhere in Europe even earlier.

Titanium patch repair to 787
 

I haven't seen this previously reported, but around the time that the ETH 787 repair was starting at Heathrow, a mile or so away in the BA hangars the 787 that was hit by a truck at Toronto last week was undergoing repair, reportedly involving a titanium patch applied to the damaged baggage door.

DavereidUK
 

So a technically poor repair has been carried out to enable the aircraft to return to service quickly, this is not unusual and the reason an aircrafts weight will grow during years of service.

The big question is if BA considers the extra weight and the disturbance of aerodynamic profile an issue that requires attention at the next major maintenance check.

Er! the repair to the BA a/c was no more than the replacement of a Ti edge protector strip to the forward cargo door. Unfortunately it had to come from the manufacturer, (in this case China) so it took over 24 hours by the time it had been shipped and gone through customs etc. These sacrificial edge protectors are in segments around each door aperature.

Thanks for the clarification, GP. It did seem rather strange that a permanent metallic repair would be made to a door, with the associated weight and aerodynamic penalty, rather than simply replacing it. That the repair was to the protective strip around the aperture makes much more sense.

Hi
Does anyone have any further info on this repair to the BA.787 like the reg of the effected airframe and which door had to be replaced/repaired?

Having assisted with a repair to a glass fibre glider fuselage, it is worth noting that it considerable effort was taken to make sure each layer of glass fibre matched that in the original fuselage in both type density and weave of glass clothas well as the direction of the fibres. The repairer in this case was a graduate of one of the German Akaflieg universities where they tend to know their stuff about composites.

The technique used was to mark both the material to be removed and the remaining structure to ensure correct alignment of the removed portion, then to burn off the resin on the edges of the removed portion to expose the fibres in the glass layers to ascertain the fibres' orientation.

Longevity of global supply chain?
 

Made me wonder how Boeing plans to make replacements of all the bits and pieces available later in life. Did they in every case obtain the rights and capability to build replacements themselves if made necessary by the demise or disinterest of the OEM?

You're holding the sheet upside down. It is Boeing who has all the rights, and licenses manufacturing to various suppliers. Subject to terms of agreement they can withdraw that license and give it to someone else or do the job themselves. Been like that since the times of the 247, nothing new.

Stop Smoking
 

The article says that for the repair, Boeing fabricated a complete new fuselage barrel, and then cut it to pieces to make a new patch. That just seems so wrong.

I trust they already looked at it in lots of detail, but I would think the smarter repair would be:

• Cut out a few test pieces from the burnt section. Test them to see how much strength was lost (and I expect it would really not be that much).
• Make a temporary patch to bring up the strength to the necessary level (if needed at all).
• Fly an unpressurized flight back to the factory (to reduce the stresses on the body and patch). Can you imagine if Boeing was able to fly the airplane back to the factory (even if unpressurized) with little or no repair work done first? That would really impress people with the strength of Boeing's "plastic body".
• Replace the whole barrel in the factory.

Maybe the above method would cost a little more, but you would have an airplane just as good as brand new. Just as strong, with no weight penalty, and no risk of any future "de-patching" problems.

But now, with the "patch" everyone will be looking at this airplane as "damaged goods" and just waiting for the top to pop off at altitude.

I think this airplane is a great candidate for a new paint job and a creative advertising company to market their stop-smoking products (patch).

No, it isn't wrong. As they said the fabricated piece was cut up to be used for other potential future repairs.
Replacing the whole barrel was looked at and it was deemed way too costly and risky due to cutting/reconnecting countless lines that run through it.

Speedbump59
 

All you have shown is how little you understand about load transfer within a composite structure and how these structures can be repaired with almost no increase in weight.

To call the repair a patch is incorrect what Boeing are doing is inserting new structure into the hole and bonding it into place using the scarf technique.

The only people who would see this sort of repair as damaged goods are those who know nothing about the subject.

Please read the posts by Volume & Barit1, these guys have a true understanding of the subject.

Not so sure...given that an aluminium aircraft may have been written off by the same incident, I think this one will stand Boeing in good stead. They couldn't have had a better demonstration that they have the ability to repair a major composite damage in the field.

I think it is a mixed bag for Boeing. Although the cause of the fire seems not to have been Boeing's fault, the bad publicity has cast further public and industry doubts on the 787.

A good repair will be great news, and increase confidence in the chosen material, but the length of time and complexity of the repair operation, plus its costs, including airport fees and lost revenue, is going to be high.

I eagerly await progress eagerly.

I think there is a lot of hype going about, questioning the strength of any patch repair. It is not beyond reason that Boing could take the opportunity to make that hole into an escape hatch, in case of any water landings. Or even a celestial observation dome for those passengers who may be amateur astronomers.
There are many other openings into the fuselage, that just have hinges and locks as the strength bearing mechanism.