Only thing that puzzles me is why cut up the remainder at all.
Seems that keeping it as whole unit would allow for a better chance of having the required segment available for random shaped damage than trying to predict in advance where to cut.

Could be based on cost to store the full article vs sections or possibly they are only cutting it into into a couple of easier to store pieces based on "cant cut here anyway" lines.

Perhaps I misspoke and there was no cutting mentioned in this news segment, possibly it only said that the whole barrel will be kept in inventory with future fixes in mind.

Thanks makes sense.
Would not be surprised if some of the news reports did say they would cut it up in advance of need.

Any time I have seen news reports on something I have first hand knowledge of I am amazed at how creative reporters can be.

"Creative" is a lot more polite than I would have phrased it. Anytime you get reporters reporting on a subject that they know little about (which appears to be most of the time), the results tend to be somewhat sketchy.

The first airline to experience a tail strike will be more than happy to buy the lower portion of that barrel...

You are again metal minded (nothing bad, that´s what 99% of the aviation people are). If you "cut out" a test piece it will fail where you clamp it into your text fixture. Composite test pieces must be especially designed and build to really load them realistically and get test results you can work with. Coupon testing of composites just makes sense for certain specific properties, otherwise you should test larger pieces of substructure.

That would require a design of the joints to allow for re-drilling and using oversized fasteners. I am not sure whether Boeing has designed the barrels for that. The fact that they had to remove the window at the joint seems to indicate that it is just barely possible to do the joint with no strength penalty. Re-doing the joint by re-drilling the holes may not be possible. Just like you can not do a simple rivet joined in the pocketed area of a chemically milled fuselage skin, the composite layup needs to have higher wallthickness and different layup for bearing strength in the area where you want to install fasteners. Unless you build your aircraft from "black metal" and waste a lot of weight on panels which are strengthened for fasteners everywhere. As we did in the past, just to allow riveted patch repair everywhere.

If Boeing is doing the composite repair right, using the same technique that is used in small composite aircraft since 1956, and regularly inspects this repair during the remaining life of that aircraft to collect data, potentially doing a destructive test of that repair once the aircraft retires, then this incident and the according repair may mean a quantum leap in large aircraft composite design. It would mean we finally build and repair large composite aircraft, not large aircraft from "black aluminium" anymore. This could be the long term test that should have been done 20 years ago to get confident in real composite repair.
In 20 years we will laugh about riveted titanium patch repair of composite aircraft, just like the small aircraft industry is laughing about what "the big guys" do up to now.
And maybe some day we will even learn that there are more clever ways to build composite aircraft than using stringer stiffened shells and ribs/frames with mouseholes cut into them joined by clips, cleats and fasteners...

B787 Dreamliner
 

Everything seems to have gone quiet on the in-service performance of this aircraft. That could be because Boeing have fixed all the problems, although that seems unlikely as some of the well-publicised problems are not going to be quick fixes.
The airworthiness authorities are presumably happy with operations continuing, so what is the airlines experience?
Apparently the flight crews love it, the engineers are not so happy as a result of almost constant modifications, and the travelling public are happy if their intended flight on a 787 occurs at all.
Sounds a win win to me.

I work on them and I am very happy...

Probably a gross exaggeration. Yes, some tweaking, etc to improve reliability is to be expected with any aircraft within first years in operation, 777 was going through the same phase.

Boeing is quoting 97% dispatch reliability - which isn't great but also isn't horrible for a new airplane (I remember one operator at 80% for the 747-400 a year after EIS :eek:). I'm no longer hearing the claim that the 787 is better than the 777 was at the same point after EIS so I'm guessing it's not :rolleyes:.

Problem areas appear to be electrical power and APU.

Not sure if this is being reported but seems to be at least twice a week in the news with this carrier

Snags soar in Dreamliner but Air India gets 10th aircraft | Firstpost

However who ever has flown it says it is a beauty....

Yep and when they finally bed down the niggling issues she'll be even better.

Can't beat a Boeing. :ok:

Recent composites conference, talk by Learjet/BBD about the Learjet85 development and certification - synopsis at Bombardier sheds light on Learjet 85 composites manufacturing : CompositesWorld

quote of relevance "Harter related that after processes were optimized, Bombardier faced the daunting task of achieving FAA certification. Much like the Boeing 787 and the Airbus A350 XWB, intensive use of composites on the Learjet 85 meant Bombardier had to perform extensive testing to satisfy the FAA’s special conditions for certification. Much of this, Harter said, focused on in-flight flammability, post-crash flammability, crashworthiness, durability, toxicity in burn, damage tolerance and thermal expansion at interactions with metals. Results, across the board, were positive. In fact, noted Harter, composite materials on the Learjet 85 outperformed aluminum in flammability and crashworthiness tests — a fact that he believes needs to be emphasized more by the aerospace composites professionals. Composite readiness testing is now complete and certification testing has begun."

Even if you discount for spin, interesting statement.

Also, with apologies for long delay in posting the following comment, I was at a major conference in the summer, where I had a chance to speak to two key (composites flammability) academic experts quoted by Amicus earlier in the thread. I posed them the question "two airplanes side by side, one aluminum, one composite, both on fire, which one would you rather be in?" One answered they would rather be in the composite airplane, the other said it made little difference.

If the replaced area of composite is going to be relatively large (more than a couple of square meters or so) I will be quite amazed if they can just brush a bit of resin on a scarfed joint and call it good to go.

Even if it is scarfed, since there will be no carbon fibers running through the glued joint, it seems difficult to believe that just a resin bond will stand up to 20 or more years of flexing and pressure cycles. I would think they would need the composite equivalent of a doubler plate across the joint to be sure.

I don't remember exactly why, but I seem to remember that for some of the very early build 787's weren't some of the fuselages disassembled and reassembled (maybe even more than once)? In that case, it seems clear that the fuselage barrels can be replaced if necessary. It may mean some new connectors in the electrical cables and hydraulic lines, but there are already plenty of those on the airplane.

Of course Boeing is very keen to show the world that it is not too difficult to repair even a large area of composite. But I really hope that they are not going to regret their decision to patch the airplane instead of replace the whole barrel.

Speedbump59

They say when in a hole stop digging !

A scarf joint replicates the way the original structure transmits the load, that is why it as strong as the original structure, the fact that Boeing lay up the structure with a wound filament is all about efficient construction rather than a need from a structural point of view and in reality each layer of the wound filament could be veiwed as one very long scarf joint.

Until you rid yourself of you metalcentric way of viewing structures you will never understand composites.

It will not be 'scarfed' in.;)

newvisitor - excellent post, too bad it wasn't here months ago when the 'composite hysteria' was in full swing.

A and C:

A scarfed joint is simply a slightly improved butt joint. With no fibers going through the joint, it does not have the same strength characteristics as the surrounding material.

Your concept of an "infinitely scarfed joint" is very nice! However it precisely highlights the reason that a scarf joint would not be used: the weakest direction in a composite is in the direction that there are no fibers, i.e delamination.

In my opinion, a simple scarf joint cannot be used for this repair. There will need to be a significant lap joint most likely with some kind of reinforcement (either interally bonded to the composite or something similar to a riveted doubler plate).

The reason for this is that for a large area of repair, strains become more critical than stresses. And a simple scarfed joint will not be able to tolerate those strains (flexing due to flight and pressure cycles) for any significant time period.

I will be happy if I am proven wrong and learn something new, but for now this is my story and I am sticking to it.

I hope that we will soon learn the details of the actual repair method so that one of us will be eating crow for dinner!

Disclaimer: the above comments have not been reviewed for accuracy by the Society for Metalcentric Thinking.

http://www.annular.org.uk/wp-content...r-joint_hr.jpg


how it is done with wood.

Speedbump59
 

I am guessing that you have never been involved In composite repair because you simply lack understanding of the subject, the company that I am involved with has been involved with the maintenance and repair of aeronautical composite structures since the late 1960's and myself have been involved since the early 90's. I am the guy who puts his name to the release to service paperwork......... We have yet to have a structural problem with an aircraft we have repaired.

Unlike a metal repair were you aim to have the edges of the repair very smooth to avoid inducing a fatigue crack the surfaces of a scarf joint are comparably rough and I am sure that under a microscope the fibers would look fluffy. This results in both sides of the repair being slightly porus, this allows the resin to soak into both sides of the repair and the fibers to intermingle within the joint in the same was as within the original structure.


As I have said before all critical structural repairs are subject to samples being tested to distruction by the original manufactures, these samples usually fail at 1-3 % above the average fail point for new items........ We have never had a test sample fail to meet the new specification.

The fact of the matter is that Boeing would never have built an aircraft structure using a totaly new Technology, composite structure has been about in aviation for 50 years now and bonding Technology since the Wright brothers, metal bonding since the DH Comet and Dove so the techniques for repair are well understood by all except the metalcentrics.

I have no idea what Boeing are planning for this repair but if it is not a scaf respair the wieght will go up for little structural advantage,

Sorry Crippen, thats a finger joint not a scarf joint. They are completely different.

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.