Cracks found in A380 wing ribs
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How the Wing is build :
I am a non pilot and a non engineer.
Would it be fair to say that the worst case scenario is that:
Would it be fair to say that the worst case scenario is that:
- time consuming and expensive checks will need to be carried out
- expensive repairs will be required
- these will add weight
- this will involve the manufacturer suffering financial penalties from customers
Does anybody have a best case scenario as well?
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Ok.The movie was deleted.We have experts around to proceed on a more professional level.
"Not much of a sheetmetal guy ..."
"Not much of a sheetmetal guy ..."
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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?
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?
Please clarify if you would, thanks,
TD
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.
a. The 7449 alloy was either off chemistry or improperly processed during subsequent heat treatments?
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?
Please clarify if you would, thanks,
TD
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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.
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Thread Starter
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update
here is an update to the thrade from the sydney morning herald rampman
More A380 wing cracks found: sources
rampmanMore A380 wing cracks found: sources
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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
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.
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:
Justin Dubon, an Airbus spokesman, said the company had established an inspection and repair procedure in coordination with the safety agency to address the cracking, which he said was linked to unforeseen stresses placed on the wings during the manufacturing process and not to a design problem.
“Both types of cracks have been traced to the manufacturing process, and those practices have now been changed,” Mr. Dubon said. All of the A380s’ wings are built at an Airbus plant in Broughton, Wales.
The first cracks were found late last year on the wing of a Qantas A380 that was being refurbished after experiencing a spectacular midair engine explosion in 2010. Airbus deemed the cracks — which extend from a bolt hole in the bracket — to be “noncritical” and advised airlines to inspect and replace the parts during routine scheduled four-year maintenance checks. The first A380s entered service four years ago.
“Both types of cracks have been traced to the manufacturing process, and those practices have now been changed,” Mr. Dubon said. All of the A380s’ wings are built at an Airbus plant in Broughton, Wales.
The first cracks were found late last year on the wing of a Qantas A380 that was being refurbished after experiencing a spectacular midair engine explosion in 2010. Airbus deemed the cracks — which extend from a bolt hole in the bracket — to be “noncritical” and advised airlines to inspect and replace the parts during routine scheduled four-year maintenance checks. The first A380s entered service four years ago.
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
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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.
"We found it is very random. We actually found them on ribs across the wing and from one aircraft to another it can be a different rib foot."
"On some aircraft we found almost none and on others we found several across the wing on both sides."
"On some aircraft we found almost none and on others we found several across the wing on both sides."
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the assembly process does have something to do with this, more so than thinking the ribs are under-designed
The new cracking is slightly different and viewed as more significant, though. It has been seen on two A380s that were being scrutinized. The damage was spotted when a customer aircraft underwent a C check within the last few weeks. One of the brackets—on its vertical part—had a more significant crack than the original hairline stresses, prompting Airbus to notify safety authorities and to launch a wider inspection of nine aircraft; evidence of cracking was found on a second aircraft.
As part of the root cause analysis, Airbus instrumented one of its own aircraft to assess whether the company’s original wing load estimates were faulty. It was determined that was not the case. The likely cause was found in the assembly process, in which too much stress is applied to the bracket when the wing skin is attached to the rib. The part itself is not being redesigned, but the assembly process is being changed as a long-term solution.
As part of the root cause analysis, Airbus instrumented one of its own aircraft to assess whether the company’s original wing load estimates were faulty. It was determined that was not the case. The likely cause was found in the assembly process, in which too much stress is applied to the bracket when the wing skin is attached to the rib. The part itself is not being redesigned, but the assembly process is being changed as a long-term solution.
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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.
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.
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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
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.
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
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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.
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
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.
Last edited by grounded27; 25th Jan 2012 at 20:24.
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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.
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!
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.
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!
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Some additional information from Airbus as reported by Flightglobal.
The airframer said the choice of alloy - designated 7449 - combined with a fastener interference-fitting process appeared to be generating the first type of crack in the feet. But a second type of crack - which EASA described as "more significant" - was also being created during the pull-down of wing skins, in the area of a butt-strap joint used between different lower skin panels.
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.
Airbus executive vice-president for programmes Tom Williams said the interim fix being carried out on affected aircraft naturally relieved these stresses, and eliminated the problem.
"We have enough ribs and feet [to conduct the repairs]," he said. Williams added that a permanent solution would look at changing the alloy - although this would require thicker rib-feet and add some 89kg in weight - and amending the pull-down process.
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.
Airbus executive vice-president for programmes Tom Williams said the interim fix being carried out on affected aircraft naturally relieved these stresses, and eliminated the problem.
"We have enough ribs and feet [to conduct the repairs]," he said. Williams added that a permanent solution would look at changing the alloy - although this would require thicker rib-feet and add some 89kg in weight - and amending the pull-down process.
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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?
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").
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").
