Well, not so completely different after all. Similar in principle (parallel orientation of spliced fibres), sort of a multiple scarf joint rotated 90 degrees, but the finger joint is used on thicker stock.

I know !! Just saying like.

:ouch:

I am completely novice in the field of large aircraft composite repairs. However I fly regularly in composite aircraft ( high performance gliders) and there when a (large) repair is done it always adds weight ( unlike in repairs in aluminium or wood/fabric) . For aerodynamically fine tuned gliders , if repair is in fuselage it can affect CG, if it is in or near control surfaces it can add instability and sometime vibrations at certains speed .
Is this irrelevant in large aircraft such as the 787 ?

A and C
 

What type of scarf joint do you believe is the most probable?

Some types of scarf joints need removing of rather large areas of original materials to be proven effective and as strong as the original type. From the side they would look like very sharp dart-like objects, one on top of the other, a bit similar as in crippen's post.

Other scarf joints are more like "step" joints (sorry, non-native English writer here..) where the "steps" can be subject to stress and creep but this saves on material.

One of the most material effective scarf joints is where you cover a piece of the original material with the repair (whilst doing this "step" joint - sorry for not knowing the proper name!), this would lead to a "bump" and may affect aerodynamics. My guess is that something like this is what happened to ATC Watcher's glideplane.

Likely Engineering Paper
 

I would expect that when the repair is finished the engineers involved will generate a 'paper' for one of the aviation engineering journals. This may not satisfy the 'peanut gallery' but for those in the business it will be welcome elucidation of the repair and repair technology used. An effective repair with no limitations on subsequent operations of the repaired airframe will be extremely good news for both Boeing and Airbus.

ATC watcher / Clandestine
 

The amount of weight a repair adds to an area of structure largely depends on the skill of the guy doing the job, there will always be some extra weight but it will always be a lot less in percentage terms than a metal repair.

Scarf joints are usually stronger than the original material by a small percentage, on a wooden aircraft you always make a test sample using off cuts of the wood you have used and the glue, once the glue has cured you take a hammer to the joint, if the wood fails that is good, if the glue fails then you have to re-do the job.

Composite work is much the same as wood when it comes to scarf joints except the glue is replaced by resin that is exactly the same as the original structure so as long as you have room to get the correct amount of scarf area there is no need for for extra structure as all you are doing is replacing like with like.

The most exacting repairs are on glider flying controls were the allowable extra weight is measured in grams, depending on how well the original manufacture was will determine if the control surface has enough margin to make a repair.

It may not happen after what they had said before that the method of repair was a confidential matter between Boeing, Ethiopian and the insurance co.

Don't the FAA and its various equivalents have to agree that the repaired aircraft is airworthy?

Poorjohn
 

That type of repair would be done by Boeing employed FAA DERs (Designated Engineering Reps.) and most likely using the FAA 8110 form, which at some point is submitted to the FAA. So it is highly likely the FAA will have a lot of oversight of the repair.

Yeah, but it doesn't necessarily mean FAA will make public the method of repair. There was a similar story with the 787's lightning protection during certification, FAA had to approve it but Boeing made sure no specifics got to public domain (they even intervened to snuff out some early press leaks/drawings) claiming their system is highly proprietary.

A and C
 

I do not question the strength of composite repairs. I am sorry if I implied such a thing.

Only thing I was pondering was how the joint would look. There are several ways to join composites and I may very well have used the wrong words.

I was thinking that the sharp long angle scarf joint would be preferred (due to strength etc), but the negative with this approach would be that a lot of unharmed material would have to be removed.

Then there are scarf joints that looks like small steps and of these there is one that I suppose have been used on the aforementioned glider.

Sorry in advance if I use wrong words. I am not familiar with the technical terms in English. I really should try to find pics to explain how I mean...

Isn't it a mute point with 'fixing composite materials' that its taken this long to fathom out just what to do/how to proceed/fix the 787 - three months + from the incident and are we any nearer solving the problems?

Clarification re scarf joins.
 

This is what I mean when I say scarf join:

http://www.jimsboats.com/scarf1.gif

Obviously Boeing would use something more like length = 15 times thickness or so.

But there are many other types of scarf joins, some that looks like steps and they can be with or without "pegs" going through the material (no such join shown in this pic but often used in wood):

http://pic3.picturetrail.com/VOL12/1...7/70507451.jpg

And here is one that has small "steps" where the join overlaps and creates a thicker material over the join. Maybe this happened to the glider. I do not find it credible that Boeing would use this method on the 787 though.

http://pressurevesseltech.asmedigita...07006jpv1.jpeg

It will not be a scarfed in joint!;)

And what problems are WE supposed to be solving? :rolleyes:

olasek

I use the term 'we' royally....

Static aircraft fire.......no repair after 3 and a half months.....

or did I miss the aircraft departing Heathrow and returning to service?

If there wasn't a few problems to solve the aircraft would have flown by now wouln't you think?

Yeah, there were a few problems to solve, the whole engineering analysis took months, agreement between various parties, FAA, etc. so what's new??

So what type of joint will it be...:ugh:

Clandestine
 

Normaly if there is room on the structure you would use what you have termed a traditional scarf joint with a face exceeding 15:1.

We have a contributor on here who keeps telling us that this won't be a scarf joint, perhaps he would like to illuminate us with his reasoning ?

A scarfed joint is an improved lap joint !
The peak stress in the bond line is "slightly" improved by factors of 50 or so compared to a lap joint, and "slightly" improved by factors of 500 or so over a butt joint.
For those who understand metal:
If you do a riveted lap joint with one row of rivets it is bloody simple to calculate. (at least if you cover the excentric issue by corrective factors for your rivet strength, which is readily avaible in the MS standards or in the SRM of your aircraft). If you do a riveted lap joint with two row of rivets, it is still all easy, the load is distributed 50/50 to the two rows. However, using flush rivets the first row is still the more critical one, as the "bypass" load around the rivets lacks the additional material of the countersunk, but that is a relatively small effect.
Now it becomes tricky, if you do a 3 row rivet joint you no longer have an even load distribution, it will be an (around) 40/20/40 load distribution. If you increase the number of rivet rows, the loads of the critical rivets do not reduce further significantly. If you would do a 10 row rivet joint, the center 4 rows of rivets will be practically free of loads, while the first row of rivets will still transfer around 25% of the load.
This can be compared to a simple bonded lap joint, the edges of the bond line do carry most of the shear stress, the center area does not transfer any laods (and that is the portion perfectly protected from all environmental influences and hence the most durable portion...).
What would you do for a riveted metal joint? You would use stepped wallthickness, you would design "fingers" on the sheets to reduce cross section according to the desired load carried. This is exactly what you would do for a bonded joint, you adapt wallthickness to the amount of load you want each sheet to carry. If you adapt it fully from full thickness to zero (practically not possible), you get a constand bondline load, hence significantly reducing the peak stress at the most vulnerable edges. As a secondary effect you also reduce excentricity of the joint, reducing secondary bending stress.

Woodworkers do it that way for centuries, and it works.
If you look at an old, crashed wooden aircraft, you typically see the wood broken, not the glue joints.

Additionally you should always remember that there is nothing like a "continuous fibre", you always have some ruptures in the filaments so it is absolutely normal that indivindiual fibres do transfer their load to neibouring fibres via the resin. This creates the famous "inherent damage tolerance" of composites material. When single fibres fail locally, their load is taken over by other fibres. Already during manufacturing of the fibres you produce a lot of broken fibres, and during part manufacturing you add more of those. Additionally not all fibres are perfectly straight, perfectly parallel or perfectly tensioned, there will always be some fibres with "slack". However, load will always be transferred to the "best" fibres via the resin matrix.
In a bonded joint exactly the same happens. The only issue is to prevent the bondline from deterioration due to the environment and from secondary stresses due to misalignment, impact etc. Therefore periodic inspections of the repair may be required. Just like you would do for a riveted metal repair.