Southwest 737 lands at Yeager Airport after hole in fuselage
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Great post Blakmax,
We are entering a new era of common usage of composite materials, my greatest fear is the numerous environmental factors that influence our ability to inspect and repair composite structures with a narrow margine of error that the system may not proactivly support given the current airline/business global culture.
We are entering a new era of common usage of composite materials, my greatest fear is the numerous environmental factors that influence our ability to inspect and repair composite structures with a narrow margine of error that the system may not proactivly support given the current airline/business global culture.
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Thanks for the support
Thanks for the supporting comments cwatters and muduckace. My main concern is the certification aspects for adhesive bonded structures and repairs. I am really concerned that the current FARs do not prevent a common (the most common) cause of failure of adhesive bonded structures and yet we see more and more structural bonds being incorporated in primary and flight critical aircraft structures. Helicopter main rotor blades are almost all exclusively adhesive bonded and they are certainly flight critical.
Maybe it is time to start a new thread. This discussion is way off the original subject, but thanks for the feedback.
Regards
blakmax
Maybe it is time to start a new thread. This discussion is way off the original subject, but thanks for the feedback.
Regards
blakmax
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Maybe we need to start a new thread because this is getting off the subject... other proprietary processes are useless.
Thanks for a most cogent explanation and one of the best contributions to general knowledge I've ever read, here or anywhere else.
Anyone who has tried to patch a hull breach in a jetski or similar composite hull can attest to the fact that you can do the same exact technique on two different parts of the hull and achieve 2 different levels of success, because of the different levels of demand on the patches and the difference in what each part of the hull was exposed to in normal use.
We've had centuries of experience with sheet metal but very little by comparison dealing with composite layups.
It's going to be very interesting to see how Boeing's 7L7 fuse barrels behave initially, and how some of ABI's composite panels - in service for years - look after a couple decades of commercial (don't give a s**t) use.
One can expect "learning situations" as we gain the same level of intuitive knowledge for composites that we already have for sheet metal.
And kudos to all the other contributors here - this is one of the least acrimonious threads I've ever read on pprune, and perhaps one of the most important.
...
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We've had centuries of experience with sheet metal but very little by comparison dealing with composite layups.
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composite repairs
Firstly thanks to rottenray for the delightful comments. After many years as a young angry man, I was once told by a very wise person that "you catch more flies with honey than you do with vinegar." That's why the responses are reasonably friendly, and I hope they will continue to be.
I noted Volume's comments
Volume, I bet that absolutely none of these were bolted repairs, so why do Boeing plan to use bolted repairs for the 787 composite structure?
You are absolutely correct about the higher skill levels for bonded composite repairs. The RAAF in Australia restructured their training in 1992 to actually teach how to perform adhesive bonding (rather than the "follow the cook book" approach) and since that time our major repair facility cut the repeat-repair rate from 43% in 1992 to almost zero, saving five man years per year of effort every year since then. Yet civil training still aims at the lowest common denominator and teach to an appallingly low standard. This is stupid. It does not take much more effort to train people to be proficient composite and bonding technicians. I personally have trained even electronics technicians to perform adhesive bonds and to do the task to the best standards.
Part of the problem with the civil system is that they use the term "composites" to encompass composite materials as well as adhesive bonding, yet they are two separate and distinct technologies. Just because a technician can lay-up composite repair patches does not mean he can prepare a metal surface to bond that repair. The blurring of the boundaries also means that FAA documents such as AC 20-107A treat the certification of adhesive bond the same as they treat the certification of composite structures. Why? The design methodologies, the failure mechanisms and testing methods are dramatically different. Composites don't fail at an interface, but adhesive bonds often do. The only similarity is that the materials are sticky before they are processed.
Consider this: Current certification of adhesive bonded structures IAW AC 20-107A involves thousands of tests at coupon, element, structural detail, sub-component and component level. If as in my previous positing, the joint was designed such that the metal (or composite structure) ALWAYS failed before the bond, then the only thing measured by thousands of tests would be the strength of the metal or composite, not the bond. So why do thousands of tests?
The approach based on designing such that the bond was never the weak element in the joint would save millions of dollars in certification costs for bonded structures. A simple test program to characterise key adhesive properties (together with a change in design methodology) would be all that was required. Further testing to demonstrate bond durability is essential, but current tests required by the FARs do not actually demonstrate bond durability.
So, if we characterise the adhesive, demonstrate that the bonding process actually produces durable bonds, use a design methodology that assures the adhesive is never the critical element, train the technicians how to perform correct bonding processes and that joint or repair will NEVER fail. NEVER. Certification costs could be dramatically reduced and reliability would dramatically increase. Am I out with the faries? No, because using this approach we have had three bond failures since 1992, and in every case, technician error (or to be more precise technician short cuts) have been found as the cause of those failures.
Now we really are off the track about the pressurisation failure, but what the heck? I suggested that the Moderator could pull me into line by moving the discussion to a seperate thread, but I've had no response, so lets keep the discussion going.
Regards
blakmax
I noted Volume's comments
We have about half a century of experience with repairing composite gliders.
You are absolutely correct about the higher skill levels for bonded composite repairs. The RAAF in Australia restructured their training in 1992 to actually teach how to perform adhesive bonding (rather than the "follow the cook book" approach) and since that time our major repair facility cut the repeat-repair rate from 43% in 1992 to almost zero, saving five man years per year of effort every year since then. Yet civil training still aims at the lowest common denominator and teach to an appallingly low standard. This is stupid. It does not take much more effort to train people to be proficient composite and bonding technicians. I personally have trained even electronics technicians to perform adhesive bonds and to do the task to the best standards.
Part of the problem with the civil system is that they use the term "composites" to encompass composite materials as well as adhesive bonding, yet they are two separate and distinct technologies. Just because a technician can lay-up composite repair patches does not mean he can prepare a metal surface to bond that repair. The blurring of the boundaries also means that FAA documents such as AC 20-107A treat the certification of adhesive bond the same as they treat the certification of composite structures. Why? The design methodologies, the failure mechanisms and testing methods are dramatically different. Composites don't fail at an interface, but adhesive bonds often do. The only similarity is that the materials are sticky before they are processed.
Consider this: Current certification of adhesive bonded structures IAW AC 20-107A involves thousands of tests at coupon, element, structural detail, sub-component and component level. If as in my previous positing, the joint was designed such that the metal (or composite structure) ALWAYS failed before the bond, then the only thing measured by thousands of tests would be the strength of the metal or composite, not the bond. So why do thousands of tests?
The approach based on designing such that the bond was never the weak element in the joint would save millions of dollars in certification costs for bonded structures. A simple test program to characterise key adhesive properties (together with a change in design methodology) would be all that was required. Further testing to demonstrate bond durability is essential, but current tests required by the FARs do not actually demonstrate bond durability.
So, if we characterise the adhesive, demonstrate that the bonding process actually produces durable bonds, use a design methodology that assures the adhesive is never the critical element, train the technicians how to perform correct bonding processes and that joint or repair will NEVER fail. NEVER. Certification costs could be dramatically reduced and reliability would dramatically increase. Am I out with the faries? No, because using this approach we have had three bond failures since 1992, and in every case, technician error (or to be more precise technician short cuts) have been found as the cause of those failures.
Now we really are off the track about the pressurisation failure, but what the heck? I suggested that the Moderator could pull me into line by moving the discussion to a seperate thread, but I've had no response, so lets keep the discussion going.
Regards
blakmax
Last edited by blakmax; 29th Jul 2009 at 11:12. Reason: sorry Mod. It was a typo. I realy didn't mean to type "noderator". Honest!
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Apparently technicians are standing on the damaged area of the fuselage to access other areas for inspection.
Of course, if one accepts that it is a step point, it probably affects what the structure should look like. So maybe this isn't a bonding or vibration fatigue failure, but instead failing to design it to be walked on?
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Surely at those skin thicknesses and frame/stringer pitches the whole top of the fusleage is 'No Step' without writing it over the whole thing?
Not forgeting that 'No Step' for a 10 stoner in trainers being careful to step on rivet lines is a bit different than an 18 stoner in hob-nailed boots, sticking his boot in the middle of panels... so take the worse case and placard for saddles and planks or a cherry-picker.
Not forgeting that 'No Step' for a 10 stoner in trainers being careful to step on rivet lines is a bit different than an 18 stoner in hob-nailed boots, sticking his boot in the middle of panels... so take the worse case and placard for saddles and planks or a cherry-picker.
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No step or a poor design?
Guys
No matter how you package this with regard to the 10 stone weakling or the 20 stone average American inspector, the real problem is that the reinforcement bonded to the skin bay reinforced everywhere but the actual section transferring the load. One side of the cracked area has skin plus frame, the other side has skin plus factory-bonded doubler. The gap in between has skin only. Now Winnie the Pooh (the bear with very little brain) could work out that if there was a pressurisation-stress fatigue or over-weight inspector overload problem anywhere, it would happen there. (Even Homer would say DUH!)
The real question is, where are the similar "reinforced" areas on the aircraft where the local stiffness and thickness changes are so significant?
Was this a BIP? (Boeing Induced Problem)
Regards
blakmax
No matter how you package this with regard to the 10 stone weakling or the 20 stone average American inspector, the real problem is that the reinforcement bonded to the skin bay reinforced everywhere but the actual section transferring the load. One side of the cracked area has skin plus frame, the other side has skin plus factory-bonded doubler. The gap in between has skin only. Now Winnie the Pooh (the bear with very little brain) could work out that if there was a pressurisation-stress fatigue or over-weight inspector overload problem anywhere, it would happen there. (Even Homer would say DUH!)
The real question is, where are the similar "reinforced" areas on the aircraft where the local stiffness and thickness changes are so significant?
Was this a BIP? (Boeing Induced Problem)
Regards
blakmax
Last edited by blakmax; 31st Jul 2009 at 12:51. Reason: Damn. That bear is spelt Pooh!
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It occurred to me to wonder if this was a cabin overpressure 'pop top' that was designed to break out in the case of a severe overpressure. There is little doubt that the skin separated at the point one would expect it to separate.
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According to this article: Southwest fuselage hole prompts Boeing 737 directive
The FAA will issue an AD tomorrow (Tuesday, Sept 14), "proposing that operators of approximately 135 US-registered Boeing 737-300, -400 and -500 aircraft begin performing external non-destructive inspections of a certain area of the fuselage skin for evidence of fatigue cracks every 500 flights".
However, "operators who have installed an external doubler in the area, as specified in a previous Boeing service bulletin, will not have to perform repetitive inspections after an initial check, providing the repair meets certain other criteria spelled out in the proposed AD".
The FAA will issue an AD tomorrow (Tuesday, Sept 14), "proposing that operators of approximately 135 US-registered Boeing 737-300, -400 and -500 aircraft begin performing external non-destructive inspections of a certain area of the fuselage skin for evidence of fatigue cracks every 500 flights".
However, "operators who have installed an external doubler in the area, as specified in a previous Boeing service bulletin, will not have to perform repetitive inspections after an initial check, providing the repair meets certain other criteria spelled out in the proposed AD".
It's a little frustrating to read an old thread like this and not find any answers; a bit like a library book with the last few pages torn out!
So here are the last couple of pages:
https://www.ntsb.gov/_layouts/ntsb.a...09FA065&akey=1
https://en.wikipedia.org/wiki/Southw...es_Flight_2294
So here are the last couple of pages:
https://www.ntsb.gov/_layouts/ntsb.a...09FA065&akey=1
https://en.wikipedia.org/wiki/Southw...es_Flight_2294
It's a little frustrating to read an old thread like this and not find any answers; a bit like a library book with the last few pages torn out!
So here are the last couple of pages:
https://www.ntsb.gov/_layouts/ntsb.a...09FA065&akey=1
https://en.wikipedia.org/wiki/Southw...es_Flight_2294
So here are the last couple of pages:
https://www.ntsb.gov/_layouts/ntsb.a...09FA065&akey=1
https://en.wikipedia.org/wiki/Southw...es_Flight_2294
