What does this have to do with the failure? Nothing!!!! 7449 alloy has failed in the desigh spec.. This is fact. I want to hear what the resolution is.

Experinece is that the result will land someplace between worst case and best case.

Does anybody have a best case scenario as well?

Ok.The movie was deleted.We have experts around to proceed on a more professionell level.




"Not much of a sheetmetal guy ..."

Shame you removed it, it did actually add to this thread and showed the parts in question quite clearly. :ok:

grounded27

There is some ambiguity in my mind relative to your proclamation. Do you mean:
a. The 7449 alloy was either off chemistry or improperly processed during subsequent heat treatments?:eek:
or,
b. The rib structure made from the alloy was marginally designed to the point the capability of the alloy was exceeded by stress levels being experienced in service?:uhoh:

Please clarify if you would, thanks,

TD

The alloy AND thickness of the rib feet was insufficient for the application. I am saying that if 7449 is a good alloy for this application there was not enough for it. If it was not it is obviously the wrong material for the job. The bottom line is that Airbus was going for weight savings in this monster of an aircraft and is now paying for a poor decision.

update
 

here is an update to the thrade from the sydney morning herald :ok: rampman

More A380 wing cracks found: sources

grounded27,

Thank you for responding and answering my questions.

First, I am reposting the video of the A-380 wing assembly process as I think it is significant. Particularly, at the times of 7:24 & 8:00, stop and look at the brackets, also known as rib feet. For each wing, there are 2,000 of these. Each wing has 62 ribs, 38 are metallic and 24 are carbon composite. The wing panels, some are very long, are creep formed to the desired contour of the wing surface at the location they will be fastened to. The rib feet of the top of the structure is designed knowing it will be in compression during actual service. The rib feet on the bottom of the structure that you cannot see are designed to be in tension and I would bet they would be different, more robust. The alloy, 7449-T7651 is a new alloy that contains no chromium unlike other 7XXX series alloys or older aluminum alloys. The T7651 condition stands for an over tempering heat treatment. This is done to maximize improvement of corrosion resistance during service. This comes at some sacrifice of other properties. What is important to understand is the significance of the assembly orientation and the possible stresses put on the rib feet that would not be realized in service.

Airbus has indicated the following:

This statement does make logical sense. The first cracks found, extending out of the bolt hole were, IMHO, coming from sharp corners. In a compressive tensile state, they would not grow very fast and would not necessarily be of grain boundary nature. Their rate of growth could be predicted and a time set for when repairs could be carried out without jepordizing safety.

The other cracks are not on the feet per-se, but on the shape below the feet in between the rib structure and the feet. These are more serious and IMHO, would be grain boundary cracks that could extend quickly depending on grain boundary length. With the orientation of the feet and the wing panel during assembly, it may be possible to unevenly stress this area depending on the force being supply to press the panel to the feet verses the weight of the panel and the sequence of bolt insertion and fastening.

Since we don't know the location of the troubled rib feet, and if the troubled feet are repeatable from wing to wing, aircraft to aircraft, it is hard to know for sure. But IMHO, the assembly process does have something to do with this, more so than thinking the ribs are under-designed. We will see as more information evolves.

TD

On Jan. 6th, Charles Champion, Airbus head of engineering, stated while discussing the cracks

On Jan 20 AW&ST reported:

Hi Turbine D,

Thanks for your post,some additional infos from my side.
" Each wing has 62 ribs, 38 are metallic and 24 are carbon composite."
Your amount is correct but officially they are talking about 23 CFRP and 49 in total.The difference comes from a center spar from rib 2 to 17.So these ribs
are half ribs but counted as one.Ref video at 7.12.

" The rib feet on the bottom of the structure that you cannot see..."

This is the opposite way.The installation of the panels in the video is showing
the bottom surface.You can identify the bottom by the man holes.

more info?
 

OK, so the video clearly shows a vacuum bag and autoclave procedure. Are the ribs bonded to the skins? Are the composite ribs cocured or secondary bonded? Are the metallic ribs bonded?

Grain boundary cracks are usually stress corrosion cracks. The clarifying evidence comes from the direction of cracking. Any designer worth his salt will always align the major loads with the rolling (L) direction, and fatigue cracks will grow perpendicular to the major loads, so they will NOT grow along grain boundaries, they will grow perpendicular to the grain boundaries. In contrast, stress corrosion cracks grow along grain boundaries so they grow parallel to the rolling direction and will be parallel to the span direction.

Stress corrosion cracking require three things: (1) an extruded, forged or rolled alloy which is susceptible to grain boundary corrosion, (2) a corrosive environment (and this may be as mild as the presence of water and some ions) and more importantly the presence of a residual stress such as would result from inadequate shimming or poor fit-up combined with the use of fasteners.

I have seen some STUPID repairs for SCC. The standard issue engineer looks at the SRM and finds "a repair for cracks in this area" and implements that repair without taking due cognisance of the crack direction. Most SRM repairs ASSUME cracks are fatigue cracks. In one case the repair required numerous fasteners to be installed ahead of the crack to be repaired, and the repair did absolutely nothing to provide restraint of crack opening.

I have been involved in a number of repair scenarios for SCC using bonded composite patches and these have been very effective. I just hope that the Airbus solution does not involve punching hundreds of fasteners through the structure to relieve local stresses due to flight loads, because if it is SCC, such repairs will be totally ineffective.

Regards

Blakmax

Turbine D,

I suppose no one could pre conceive the results of in service stresses, and I suppose this is an expected factor ...To an extent. From the alloy's that were used to general design of the aircraft, much is new. It is certain that more problems will reveal themselves over time (it does happen with all aircraft), it is the structural ones like this that have potential to cost $$$$$ in loss of time in service.

I would be just as critical of Boeing's composite fuselage. We are engineering aircraft w/o set precedent. We are using process standards based on what we know, skating the razors edge of the unknown.

Yes, of an aluminum wing skin panel being shaped & treated post milling.

Hi no-hoper,

Thanks for the information and correction of the wing manufacturing video. I think you are correct relative to the panel being placed against the ribs, it would appear to be an underside of the wing panel.

blakmax,

What you are seeing wrapped in the plastic bag is the wing panel including the fixture it is resting on. The fixture defines the desired shape (contour) that will
be established by the "baking process". After removal from the autoclave the wing panel is removed from the fixture. The skin is not bonded to the ribs per-se, but bolted using fasteners. For the composite ribs, a special fastner or bolt was designed by Alcoa working with Airbus for fastening metal to composite, see a picture of it below which was contained in an magazine article about Alcoa.
http://i1166.photobucket.com/albums/...2/lockbolt.jpg

I am not sure the slabs or thick plates, from which the ribs are machined, are rolled at all. If they are not, the grains would have random orientations based upon the solidification pattern. Repair of the cracks, assuming they are readily assessable, might be acomplished by welding as the alloy, IMO, is weldable by TIG or Stir welding processes.

Machaca,

Thanks for all the information you have provided. Airbus' Charles Champion's remarks concerning the dispersion of cracking makes for an interesting search process I would think!

"Australian Licensed Aircraft Engineers Association"

The usual problem in English.... confusing 'mechanics' and 'engineers'.

And confusing "association" and "union".

Some additional information from Airbus as reported by Flightglobal. :uhoh:

What's the worry?
Anybody ever seen toilet paper tear properly on the perforations?

More seriously, a question.
Are there any pictures (photos, drawings, diagrams) that could tell this ancient a bit more in detail where exactly the problem occurs?

If I'm reading and viewing all this right, the likely failure would be a wing skin panel tearing free from the ribs at the attachment point (the little dog-biscuit-shaped "feet"). Correct?

Which wouldn't be good (lot of noise, lot of drag, loss of a percentage of lift, possible jamming of loose panel in flight controls) - but is a lot different from the ribs or spars themselves departing the airframe (the wing "falling off").

ChristiaanJ,

The best photo of the ribs containing the wing brackets or wing rib feet is depicted in machaca's post #78 on page #3, first photo. The multiple rectangular things that stick upward out of each rib, each with four holes, are the feet. I assume the type 1 cracks come out of one or more of these holes in certain locations. The shape that connects these to the actual rib is where I think the more serious type 2 cracks developed. I guess the cracks are more random in nature that what I would have thought.

pattern_is_full,

You are correct, that is the way I would read it as well.

Turbine D

Every landing is different. I was turned on by a a hard landing at Osh Kosh recorded below.


This shows the enormous wing flex during a hard landing with a walk through. Every landing is different as I would expect every crack to be (hard or not). Be it a hard landing or not, be it a large crack or not in these rib feet. Sideload/torque in a crosswind is probably a significant factor to add to the random nature.

[quote]If I'm reading and viewing all this right, the likely failure would be a wing skin panel tearing free from the ribs at the attachment point (the little dog-biscuit-shaped "feet"). Correct? [unquote]

Don't think so. The likely failure would be that one bolt (out of over 100) would not be transmitting torsion from skin into that one rib. The ribs don't carry any bending loads, just torsion and curvature crushing loads between top and bottom skins.

My 2 cents.
The rib feet cracking is in both carbon and alloy ribs.
The cracking has been found across the fleet (all airlines)
Some cracking has been found on wings not yet fitted.
The root cause seems to be in manufacture.
The fit of the fastener (similar to pictured in Turbine D post) has to much negative tolerance in the fit and therefore induces the cracks some of which can be seen by the naked eye others are found by NDI.
The inspection /repair sched is for the C2 check which will probably add many extra days to the ground time

This info is not knew and I am suprised it has taken this long to hit the prune :ok:

According to this ;

;aluMATTER*|*Aluminium*|*Wrought Aluminium Alloys*|*Examples of Applications

7449 is a wrought alloy.

Reading the assembly description, it seems that this is a new process ;
Higher Levels of Automation Lift Productivity for Airbus A380 Wing Assembly Process

" For the Airbus A380 panel-production facility, Electroimpact built four machine lines, each with two machines for upper and lower surface panels. Each line includes three fixtures, where four panels are loaded. The jigs hold the components in accurate form and location while the automated machines drill, rivet, and bolt the components together. Sealant is applied to the components during the jig load. No temporary fasteners are used.
Thus, after fastening, the wing panel assemblies are complete. No interim operations are needed to clean and deburr. The one-up assembly process reduces handling damage and positioning inaccuracies (datum errors). The machines can install rivets and bolts in diameters of ¼ to ½ in., with a stack range up to 2.5 in. Automated cold working, hole probing, countersink sealing, and collar installation are all included. "


[quote]
" In the region of rib 26 and stringer 21, larger-than-expected gaps - some 1.5-2mm rather than 0.5mm - between the sections involved in the pull-down had resulted in stresses being induced, leading eventually to cracking under the wear of normal airline operations. " [quote]

Are the hole tolerances such that with a 2.0 mm gap lateral forces from fastener misalignment could also be introduced ?

User airborne on aero.de - Luftfahrt-Nachrichten und -Community found this link.Thanks !

http://www.doricassetfinance.com/pdf...pdate_a380.pdf

A380 Wing Cracks
 

Apologies if posted elsewhere...

The BBC are reporting that the A380 fleet requires inspections...

BBC News - Airbus to inspect all A380 superjumbos for wing cracks

I hear that the EK boys on the A380 are all on standby from next month? any one out there any info on this?

I always said I was glad I fly Boeing as I used to be a car dealer and French cars were shockingly bad! seems the planes maybe going the same way :}

Isn't it rather a tenuous link between French cars and aircraft wings manufactured in Wales or Chester or Filton or wherever?

Funny how they've not picked up on the latest 787 problem -
Boeing Orders Checks of 787 Dreamliners on Fuselage Delamination - Businessweek

herotozero
 

you are right about French cars...but Airbus is not French.
It is European !
and the wings are made in the UK.
and whilst I am not a lover of things"French", I think Boeing have just as many problems with new aircraft. .
Think B787!!
and compare delays/problems with A380
It just silly to blame the French.

"It just silly to blame the French."


You don't know how the internet works do you? :E

Did someone not enter the right information into the stress analysis model? Very unusual to have any cracking so early in service life (anywhere!). I thought tombstone design went out with the DC10.

From Flight Global:

Looks like the method of assembly is actually causing the problems.

A380 wing problems.
 

And the wings are made in the U.K.--Hmmmmmmmmm!.

this on Reuters

Europe to extend Airbus A380 checks - sources
Wed, 08/02/2012 - 10:21
By Tim Hepher

KUALA LUMPUR (Reuters) - European air safety officials are preparing to extend checks for Airbus A380 wing cracks to the entire superjumbo fleet, sources close to the matter told Reuters on Wednesday.

The move to inspect all 68 A380s in service came as Qantas Airways grounded one of its planes, saying engineers had found 36 wing cracks after the aircraft encountered severe turbulence.

By signalling that the defects may be structural and widespread, the fleet-wide inspection order will refocus attention on flaws identified in flagship jets at both Airbus and Boeing . The aircraft makers maintain that their newest jets remain safe to fly after problems were caught at an early stage.

"This is an extension of a process already underway," said one of the people, who asked not to be named. "An effective repair has been identified."

Airbus, a unit of European Aeronautic Defence & Space Co. , declined to comment on the additional inspections. A spokesman for the European Aviation Safety Agency (EASA) was not immediately available.

The aviation watchdog last month ordered checks on one-third of the A380 fleet after cracks were found in a handful of the thousands of L-shaped brackets that fix each wing's exterior to its internal ribcage-like structure.

EASA has yet to set out a timetable for the new inspections, two aviation sources said. Planes will be checked as they cross wear-and-tear thresholds at which the tiny cracks become detectable.

Inspectors had initially focused on 20 aircraft operated by Singapore Airlines , Air France and Dubai's Emirates - which have logged the most A380 flights since the double-decker plane entered service four years ago.

(Reporting by Tim Hepher; writing by Laurence Frost; Editing by Geert De Clercq and Jane Merriman)

French? British?


From Machaca’s post 78 referred to above, the machine that does the fixing appears to be made by: “ . . . Electroimpact Inc., Mukilteo, Wash., the prime contractor for wing-assembly automation tools.”


Now which country is that?


IMHO – the country is irrelevant. The problem will be resolved, as engineering issues are in well regulated companies, by the usual disciplines, not by racist or other xenophobic point scoring. Nor by unqualified SLF opining freely.


Chris N

Perhaps this thread has been hijacked, I don't know. What I do know however is regardless of whether this gets sorted now or not and I genuinely believe it will, that isn't actually the issue.

We should be asking what sort of regulatory oversight is in place thats first denies an issue, then stalls on the inspection process for the rest of the fleet and then finally admits to an issue that engineers had been stating was there from day one.

The time scale is unacceptable and so don't be surprised if conspiracy theorists start suggesting cover ups or inventing links between Airbus the european manufacturer and Easa the european regulator.

I personally see one issue here and only one. I can no longer distinguish between regulator and operator. To me they are one and the same. Thats dangerous in my view.

The regulatory issue is called Continued Airworthiness and has its own part under the codes.

It presumes that some degradation may occur over the life of the fleet and that the type holder/operator must provide an analysis and a program to address the issue in a manner that minimizes (not eliminate) the risk over the time period that it is in operation.

If the risk is high (compared to all other risks) than the time period that it is allowed to exist will be short. The implication of this that a fleet is always under a degree of risk for all other problems (known and unknown) and that no single known problem du jour, should significantly contribute

My read of the current discussion on this problem is that it will not significantly contribute to overall risk in the near time-frame, but must be addressed before accumulating even more risk (wear out mode).

The type holder, has apparently proposed a corrective action and the regulator accepted this with the understanding that as new data is found the corrective action program will be updated.

It's a validated process and I really don't see how we on the outside can pick it apart without new data