Southwest FLT 812 Decompression and diversion
Boeing lap joints and bonding
I have no inside knowledge of the SWA failure but was heavily involved in corrective action after Aloha accident; see [/FONT]http://avstop.com/stories/aloha.html[FONT=Arial] for that story. So I'll make an informed guess of what happened.
Hopefully the problem is confined to Boeing 737s up to LN 288, because these used a bonding process that was found to be defective; the process was relied upon to reduce fatigue stresses across longitudinal lap joints in the fuselage. Aloha happened after the bond failed and skin cracked at the top row of rivets; which were critical because they were countersunk. Fix was to replace all those rivets with button heads. BTW you can spot an old, suspect aircraft by the prominent rivets. This did nothing to aleviate criticality of bottom row. Just postponed the evil day and my understanding is that it is the bottom row that failed on SWA.
For aircraft structures, bonding has special attraction of minimizing stress concentrations and associated fatigue failures that seem inevitable with rivets or welding. [/COLOR][COLOR=#000000]However adhesive bonding is never as easy as it seems. A major reason is difficulty in establishing, substantiating and maintaining process specifications that ensure bond reliability and durability. Not only is the adhesive prone to variability depending on temperature cycles, pressure etc. Perhaps more so, durability depends on surface preparation.
Boeing learnt these lessons the hard way, with a series of embarrassing failures. The worst was the Aloha B737 fuselage failure in 1988. In (Australian) Bureau of Air Safety Investigation Journal, June 1992 I wrote:
“(The fuselage) is made in panels which are typically about 4 metres long and 2 metres wide. These are joined together with rows of rivets. Obviously the skin is weaker where it is drilled for the rivets, so on early B737's Boeing engineers tried to reinforce the joints with epoxy adhesive. It was these joints which failed first and let the skins rip away from the aircraft.
The first Boeing 737 was delivered in 1967, the Aloha aircraft was delivered in May 1969 and by about then Boeing became aware of problems with the adhesive bonding process. The adhesive worked much like a two tube mix used by a home handyman except that the glue was premixed and held on a scrim tape. By keeping it refrigerated the adhesive reaction was suspended. At the right time in construction the adhesive tape was laid between the skins, these were riveted together and the glue then cured as it warmed up to room temperature. That was the theory. In practice the adhesive did not really bond to the aluminium skin, it only bonded to the very thin layer of oxide on the surface of the aluminium. Attachment of the oxide film to the metal underneath was dangerously variable. Also if the scrim was too cold when it was applied it attracted condensation which prevented proper adhesion. If the scrim got too warm it partially cured before it was in place and again adhesion failed. Whenever adhesion failed the rivets and surrounding skin were overloaded and the skin began to crack.
Boeing was not secretive about the problem. The bonding deficiencies and their rectification were discussed in many technical papers in the early 1970's. The whole U.S. industry was embarrassed because the Europeans had been successfully bonding aircraft for 30 years. In 1975 the U.S.A.F. stepped in with a large contract to catch up with the Europeans. It went to Boeing's arch rival, Douglas.
Boeing progressively improved the design of the skin joints and hoped that for aircraft already in service the problem could be controlled with enhanced inspections. From May 1970 onwards Boeing sent the airlines a series of bulletins recommending inspection and sealing of the joints. Cracking and corrosion still went on. Diligent airlines found it and fixed it. Less diligent ones stayed lucky.”Of the tech papers mentioned above specific mention needs be made of “Surface Preparation – the key to bond durability” by Boeing Chief Chemist Corey McMillan. It spells out the evolution of surface preparation and failures along the way.
Problem is that many aircraft are still flying which were built using surface preparations of dubious repute!
Apart from the faulty cold bond technique that led to Aloha, we knew at the time of Aloha that Boeings hot bond technique was better but not completely reliable. A year earlier in 1987 hot bond failures on B747 were addressed by SB 747-53A2279, FAA AD 87-16-13 and Australian AD/B747/59.
At the time, Australia tried but never really got Boeing or FAA to acknowledge the risk of hot bond failures on B737; that hazard still lurks!
Hopefully the problem is confined to Boeing 737s up to LN 288, because these used a bonding process that was found to be defective; the process was relied upon to reduce fatigue stresses across longitudinal lap joints in the fuselage. Aloha happened after the bond failed and skin cracked at the top row of rivets; which were critical because they were countersunk. Fix was to replace all those rivets with button heads. BTW you can spot an old, suspect aircraft by the prominent rivets. This did nothing to aleviate criticality of bottom row. Just postponed the evil day and my understanding is that it is the bottom row that failed on SWA.
For aircraft structures, bonding has special attraction of minimizing stress concentrations and associated fatigue failures that seem inevitable with rivets or welding. [/COLOR][COLOR=#000000]However adhesive bonding is never as easy as it seems. A major reason is difficulty in establishing, substantiating and maintaining process specifications that ensure bond reliability and durability. Not only is the adhesive prone to variability depending on temperature cycles, pressure etc. Perhaps more so, durability depends on surface preparation.
Boeing learnt these lessons the hard way, with a series of embarrassing failures. The worst was the Aloha B737 fuselage failure in 1988. In (Australian) Bureau of Air Safety Investigation Journal, June 1992 I wrote:
“(The fuselage) is made in panels which are typically about 4 metres long and 2 metres wide. These are joined together with rows of rivets. Obviously the skin is weaker where it is drilled for the rivets, so on early B737's Boeing engineers tried to reinforce the joints with epoxy adhesive. It was these joints which failed first and let the skins rip away from the aircraft.
The first Boeing 737 was delivered in 1967, the Aloha aircraft was delivered in May 1969 and by about then Boeing became aware of problems with the adhesive bonding process. The adhesive worked much like a two tube mix used by a home handyman except that the glue was premixed and held on a scrim tape. By keeping it refrigerated the adhesive reaction was suspended. At the right time in construction the adhesive tape was laid between the skins, these were riveted together and the glue then cured as it warmed up to room temperature. That was the theory. In practice the adhesive did not really bond to the aluminium skin, it only bonded to the very thin layer of oxide on the surface of the aluminium. Attachment of the oxide film to the metal underneath was dangerously variable. Also if the scrim was too cold when it was applied it attracted condensation which prevented proper adhesion. If the scrim got too warm it partially cured before it was in place and again adhesion failed. Whenever adhesion failed the rivets and surrounding skin were overloaded and the skin began to crack.
Boeing was not secretive about the problem. The bonding deficiencies and their rectification were discussed in many technical papers in the early 1970's. The whole U.S. industry was embarrassed because the Europeans had been successfully bonding aircraft for 30 years. In 1975 the U.S.A.F. stepped in with a large contract to catch up with the Europeans. It went to Boeing's arch rival, Douglas.
Boeing progressively improved the design of the skin joints and hoped that for aircraft already in service the problem could be controlled with enhanced inspections. From May 1970 onwards Boeing sent the airlines a series of bulletins recommending inspection and sealing of the joints. Cracking and corrosion still went on. Diligent airlines found it and fixed it. Less diligent ones stayed lucky.”Of the tech papers mentioned above specific mention needs be made of “Surface Preparation – the key to bond durability” by Boeing Chief Chemist Corey McMillan. It spells out the evolution of surface preparation and failures along the way.
Problem is that many aircraft are still flying which were built using surface preparations of dubious repute!
Apart from the faulty cold bond technique that led to Aloha, we knew at the time of Aloha that Boeings hot bond technique was better but not completely reliable. A year earlier in 1987 hot bond failures on B747 were addressed by SB 747-53A2279, FAA AD 87-16-13 and Australian AD/B747/59.
At the time, Australia tried but never really got Boeing or FAA to acknowledge the risk of hot bond failures on B737; that hazard still lurks!
Join Date: Apr 2002
Location: FUBAR
Posts: 3,348
Likes: 0
Received 0 Likes
on
0 Posts
Ironically the problem this time has been caused by a mod to the design designed to reduce the risk caused by cracking.
Distance between tear strips was increased from 10 to 20 inches to reduce possibility of cracking propogating from frame to frame as is believed to have happened in the Aloha incident.
This redesign is believed to have led to this problem, hence why only 579 aircraft manufactured between 1993/2000, of which 175 are believed to be above the 30,000 cycle trigger are affected.
Best laid plans of mice & men
Distance between tear strips was increased from 10 to 20 inches to reduce possibility of cracking propogating from frame to frame as is believed to have happened in the Aloha incident.
This redesign is believed to have led to this problem, hence why only 579 aircraft manufactured between 1993/2000, of which 175 are believed to be above the 30,000 cycle trigger are affected.
Best laid plans of mice & men
Join Date: Jul 2008
Location: Australia
Posts: 372
Likes: 0
Received 0 Likes
on
0 Posts
Aloha bonding
Ozaub
Your assessment of Aloha 243 is close. The problem with the bonding was not the mixing process, because they did not mix the adhesive. They used a film adhesive, similar to double sided tape. The root of the problem was that the material they used was a room temperature curing system, so that they did not have to heat the structure (as the Europeans always did). That saved complex heating procedures for large structures and was much quicker.
The essence of the problem is the requirement to cure at room temperature. Because the material is a pre-mixed tape, it must be stored frozen, or the material will cure off prematurely. When the material was cut is was frozen and when it was applied to the structure it was frozen. As a consequence, atmospheric moisture condensed on the cold surface. Now adhesive bonds depend upon chemical bonds formed at the interface during the curing process. These chemical bonds give the adhesive strength and also durability. In the case of the early 737s and some 727s and 747s as well, the moisture inhibited adequate contact between the ahdesive and the metal, resulting in weak bonds which later failed in service. The disbonding led to higher than expected loads at the fasteners, which when combinded with knife-edge countersinks led to the fatigue cracking at multiple sites, and then to the failure once the small cracks linked up.
Adhesive bonded structures can be very reliable when the processes are correctly validated and implemented. In this case they were not.
Your assessment of Aloha 243 is close. The problem with the bonding was not the mixing process, because they did not mix the adhesive. They used a film adhesive, similar to double sided tape. The root of the problem was that the material they used was a room temperature curing system, so that they did not have to heat the structure (as the Europeans always did). That saved complex heating procedures for large structures and was much quicker.
The essence of the problem is the requirement to cure at room temperature. Because the material is a pre-mixed tape, it must be stored frozen, or the material will cure off prematurely. When the material was cut is was frozen and when it was applied to the structure it was frozen. As a consequence, atmospheric moisture condensed on the cold surface. Now adhesive bonds depend upon chemical bonds formed at the interface during the curing process. These chemical bonds give the adhesive strength and also durability. In the case of the early 737s and some 727s and 747s as well, the moisture inhibited adequate contact between the ahdesive and the metal, resulting in weak bonds which later failed in service. The disbonding led to higher than expected loads at the fasteners, which when combinded with knife-edge countersinks led to the fatigue cracking at multiple sites, and then to the failure once the small cracks linked up.
Adhesive bonded structures can be very reliable when the processes are correctly validated and implemented. In this case they were not.
Join Date: Mar 2008
Location: Oakland, CA
Posts: 43
Likes: 0
Received 0 Likes
on
0 Posts
If I may be permitted a daft SLF question - I only drive little airplanes with the small wheel at the back - I'm not sure really sure what correction Blakmax is making here. Ozaub described the bonding process in exactly the same way and listed condensation as one of the possible causes of the bonds failing. What'd I miss?
Regardless, I'm getting a great education here - thanks, folks.
Regardless, I'm getting a great education here - thanks, folks.
Join Date: Sep 2010
Location: earth
Posts: 1,341
Likes: 0
Received 0 Likes
on
0 Posts
Fact is despite the cycles and hrs, none of us can compile the data that resulted in this failure. Starting from the last rivits holding the skin, unknown "g" stress from hard/semi hard landings to turbulence (largest unmeasurable factor), the environment over 15 years "hot/cold/wet/dry". Just to speculate that this piece of skin of this aircraft had a poor run of luck based on the factors above is probably in atleast in one way the truth.
Cycles and hrs are not a factor given this type aircraft. Allthough the crown is least inspected and incinedence like this will likely prove for for more detailed inspection.
Human factors, pilots are often on the fence to report a (semi hard landing), this area of the aircraft is not scrutinized by maintenance and difficult to inspect at the gate. These are factors that will change.
Cycles and hrs are not a factor given this type aircraft. Allthough the crown is least inspected and incinedence like this will likely prove for for more detailed inspection.
Human factors, pilots are often on the fence to report a (semi hard landing), this area of the aircraft is not scrutinized by maintenance and difficult to inspect at the gate. These are factors that will change.
Last edited by grounded27; 7th Apr 2011 at 05:20.
Grounded:
Rubbish. Why do you think we maintain them at set time intervals and life a lot of components in terms of hours or cycles? Didn't you know that the Aloha aircraft was the cycle leader of the fleet?
Did I waste Six years of my life looking at cycles and hours flown then making judgements about component reliability?
Cycles and hrs are not a factor given this type aircraft.
Did I waste Six years of my life looking at cycles and hours flown then making judgements about component reliability?
Join Date: Apr 2002
Location: FUBAR
Posts: 3,348
Likes: 0
Received 0 Likes
on
0 Posts
readywhenreaching,
I am guessing that list is merely compiled from production line number, the affected batch were specified on the AD, but takes no account of cycles.
Of 579 aircraft in the batch, only 175 are thought to have reached the 30,000 cycle trigger point.
I am guessing that list is merely compiled from production line number, the affected batch were specified on the AD, but takes no account of cycles.
Of 579 aircraft in the batch, only 175 are thought to have reached the 30,000 cycle trigger point.
Join Date: Jul 2008
Location: Australia
Posts: 372
Likes: 0
Received 0 Likes
on
0 Posts
For martinprice. The point is that it was an active decision by the manufacturer to choose a specific form of adhesive which was highly susceptible to moisture problems. It had nothing to do with the mixing of the adhesive. It was therefore not a quality control problem, it was a design/material selection issue.
Hope that point is clearer now.
Thanks for the interest.
Hope that point is clearer now.
Thanks for the interest.
Photos in Post #126
The second piece of skin in those pictures seems to have a "hump" or bulge to the middle section of skin between the longitudinal rows of rivets.
Is that likely just due to tension when the adjoining piece tore loose? Or from the disassembly procedure? Or something else?
Is that likely just due to tension when the adjoining piece tore loose? Or from the disassembly procedure? Or something else?
The fuselage is not circular, it is ovoid. 
There are bloody great beams running across each frame at at deck level that support the floor. They are in tension when the aircraft is pressurised and keep it in its shape.
This is the B707 profile it is exactly the same as the 727 and 737.
There are bloody great beams running across each frame at at deck level that support the floor. They are in tension when the aircraft is pressurised and keep it in its shape.
This is the B707 profile it is exactly the same as the 727 and 737.
Last edited by Sunfish; 9th Apr 2011 at 01:33.
Join Date: Jan 2008
Location: 3rd rock from the Sun
Posts: 15
Likes: 0
Received 0 Likes
on
0 Posts
Sunfish
Yes I am aware of the true shape of a B737 fuselage. In much greater detail than I would care to.
I was sure when I posted my hastily written explanation of the 'bulge' in the cutout lap section that someone would correct me, and you have excelled with that.
Thanks for the detailed post.
Yes I am aware of the true shape of a B737 fuselage. In much greater detail than I would care to.
I was sure when I posted my hastily written explanation of the 'bulge' in the cutout lap section that someone would correct me, and you have excelled with that.
Thanks for the detailed post.
Join Date: Sep 2010
Location: earth
Posts: 1,341
Likes: 0
Received 0 Likes
on
0 Posts
Sunfish
Rubbish. Why do you think we maintain them at set time intervals and life a lot of components in terms of hours or cycles? Didn't you know that the Aloha aircraft was the cycle leader of the fleet?
Did I waste Six years of my life looking at cycles and hours flown then making judgements about component reliability?
Did I waste Six years of my life looking at cycles and hours flown then making judgements about component reliability?
Don't get me wrong cycles and HRS are a solid rule of measure but not accurate, not absolute.
The answer to your second question is hard to answer. Did you enjoy your job, were you compensated well, were you appreciated for your skills? I am sure you count beans very well, not to belittle your task, just saying that the bottom line is you were payed to SPECULATE based on statistics.
@ grounded spanner and sunfish
Sure, I know the aircraft skin is curved. But to my eye (trained in photography, photo analysis and graphics arts), the curvature is different in that particular piece of metal in the area between the rows of rivets, than it is overall from edge to edge.
I.E., I see, in small scale, the same kind of change in curvature radius as at the "top of floor beam" in sunfish's X-section.
(And, yes, I know the piece of metal pictured does not come from the area of the floor beam
- which is why I find the change in what should be a smooth, constant-radius rooftop curve interesting.)
In the first picture, the reflection of the overhead lights makes a sharp change in shape as it crosses the lower rivet line. Which indicates a change in the curvature at that point - a distortion of the smooth curve.
In the third picture, end-on, there is also a change in curvature visible between the rivet lines.
I'd diagram what I'm seeing, but pprune doesn't allowed for direct uploading of images.
Sure, I know the aircraft skin is curved. But to my eye (trained in photography, photo analysis and graphics arts), the curvature is different in that particular piece of metal in the area between the rows of rivets, than it is overall from edge to edge.
I.E., I see, in small scale, the same kind of change in curvature radius as at the "top of floor beam" in sunfish's X-section.
(And, yes, I know the piece of metal pictured does not come from the area of the floor beam
- which is why I find the change in what should be a smooth, constant-radius rooftop curve interesting.)In the first picture, the reflection of the overhead lights makes a sharp change in shape as it crosses the lower rivet line. Which indicates a change in the curvature at that point - a distortion of the smooth curve.
In the third picture, end-on, there is also a change in curvature visible between the rivet lines.
I'd diagram what I'm seeing, but pprune doesn't allowed for direct uploading of images.
Join Date: Jan 2008
Location: 3rd rock from the Sun
Posts: 15
Likes: 0
Received 0 Likes
on
0 Posts
pattern_is_full
Now that I look at the photos on something bigger than an iPhone, and with your detailed description, I do see what you are talking about.
The change in curvature in the first photo (something akin to a slight crease) is very normal with a removed skin section across a rivet line. It is sometimes called 'pillowing'. You can often see on an aircraft fuselage that the skin changes profile at a rivet line. This is because as the aircraft is pressurised, it stretches ever so slightly, and when unpressurised, the stringers behind (only slightly) return to a smaller diameter circle. Nothing to be concerned about as long as it is not pronounced and/or there are signs of distress (cracks, pulled rivet heads etc). That photo looks very normal (ignoring the torn section next to it!!).
Whether there is any extra curvature is impossible to tell. If there were any extra curvature in the skin section it would have to be material that had come from the lap joint. I cannot see any signs of paint distress at the lap joint, which means that the skin will not have moved by more than 1/16 to 1/8 of an inch. Any more than that and the lap would have let go anyway.
I would lay money that that unbroken section of skin contains cracks (significant ones) in the lower row of fasteners, just due to its proximity to the torn section, but that the cracks are not sufficiently large to continue to link up yet. The NTSB will be very interested in how much longer that section had before it would let go.
Now that I look at the photos on something bigger than an iPhone, and with your detailed description, I do see what you are talking about.
The change in curvature in the first photo (something akin to a slight crease) is very normal with a removed skin section across a rivet line. It is sometimes called 'pillowing'. You can often see on an aircraft fuselage that the skin changes profile at a rivet line. This is because as the aircraft is pressurised, it stretches ever so slightly, and when unpressurised, the stringers behind (only slightly) return to a smaller diameter circle. Nothing to be concerned about as long as it is not pronounced and/or there are signs of distress (cracks, pulled rivet heads etc). That photo looks very normal (ignoring the torn section next to it!!).
Whether there is any extra curvature is impossible to tell. If there were any extra curvature in the skin section it would have to be material that had come from the lap joint. I cannot see any signs of paint distress at the lap joint, which means that the skin will not have moved by more than 1/16 to 1/8 of an inch. Any more than that and the lap would have let go anyway.
I would lay money that that unbroken section of skin contains cracks (significant ones) in the lower row of fasteners, just due to its proximity to the torn section, but that the cracks are not sufficiently large to continue to link up yet. The NTSB will be very interested in how much longer that section had before it would let go.
The stupidity is coming thick and fast now.
Grounded27:
I'm an engineer not a bean counter. The reason we use cycles and hours has nothing to do with bean counting. Those measures are surrogates for the strain history of the part in question and in metals that don't have a yield point (ie Everything except steel) the stress history determines when the component will fracture.
We make estimates based on experiment as to how long a component will last, then we apply safety factors to that estimate. When the aircraft is in service, we monitor all failures and continue testing to refine those estimates based on experience. In more than one case I dealt with, we actually reduced the number of failures by extending the time in service limits for an assembly, see if you can work out why.
In other words, try telling a turbine disk or a chunk of aluminium that its life in service is really infinite, it's just that bean counters make us change them.
Ozaub has already explained that Boeing had a problem with quality control in relation to the adhesive system it once used. That does not mean that the process was defective, it means that Boeing discovered after the fact that it could not precisely control the manufacturing conditions tightly enough to ensure a sufficiently reliable bonding system.
Before you yappers now scream for more of Boeing s blood, by "sufficiently", I do not mean 100% guarantee, I mean to an extent to where any defects are small enough to not result in stress concentration likely to cause severe crack growth during the expected service life of the aircraft which is god knows how many thousand hours and cycles (60,000 hrs? 60,000 cycles?).
The bloody aircraft has not done badly considering, and the safety measures worked as advertised. The only issue for Boeing is that the problem surfaced considerably earlier than expected.
The question is now to work out how extensive the problem is and determine what the best inspection methods are and what the repair schemes are.
I'm also getting fed up with what I term "The pprune effect" whereby every self proclaimed expert offers a solution from within their own expertise.
By that I mean; when there is an aircraft accident, Prune attracts the computer expert who posits the cause as a software problem. The chemist suggests the fuel was faulty. The teacher wonders about the pilot training and the lawyer blames it all on criminal negligence by the designer.
If you have never worked with aircraft, I wish some of you might be a little more tentative in coming to conclusions.
Grounded27:
Don't get me wrong cycles and HRS are a solid rule of measure but not accurate, not absolute.
The answer to your second question is hard to answer. Did you enjoy your job, were you compensated well, were you appreciated for your skills? I am sure you count beans very well, not to belittle your task, just saying that the bottom line is you were payed to SPECULATE based on statistics.
The answer to your second question is hard to answer. Did you enjoy your job, were you compensated well, were you appreciated for your skills? I am sure you count beans very well, not to belittle your task, just saying that the bottom line is you were payed to SPECULATE based on statistics.
I'm an engineer not a bean counter. The reason we use cycles and hours has nothing to do with bean counting. Those measures are surrogates for the strain history of the part in question and in metals that don't have a yield point (ie Everything except steel) the stress history determines when the component will fracture.
We make estimates based on experiment as to how long a component will last, then we apply safety factors to that estimate. When the aircraft is in service, we monitor all failures and continue testing to refine those estimates based on experience. In more than one case I dealt with, we actually reduced the number of failures by extending the time in service limits for an assembly, see if you can work out why.
In other words, try telling a turbine disk or a chunk of aluminium that its life in service is really infinite, it's just that bean counters make us change them.
Ozaub has already explained that Boeing had a problem with quality control in relation to the adhesive system it once used. That does not mean that the process was defective, it means that Boeing discovered after the fact that it could not precisely control the manufacturing conditions tightly enough to ensure a sufficiently reliable bonding system.
Before you yappers now scream for more of Boeing s blood, by "sufficiently", I do not mean 100% guarantee, I mean to an extent to where any defects are small enough to not result in stress concentration likely to cause severe crack growth during the expected service life of the aircraft which is god knows how many thousand hours and cycles (60,000 hrs? 60,000 cycles?).
The bloody aircraft has not done badly considering, and the safety measures worked as advertised. The only issue for Boeing is that the problem surfaced considerably earlier than expected.
The question is now to work out how extensive the problem is and determine what the best inspection methods are and what the repair schemes are.
I'm also getting fed up with what I term "The pprune effect" whereby every self proclaimed expert offers a solution from within their own expertise.
By that I mean; when there is an aircraft accident, Prune attracts the computer expert who posits the cause as a software problem. The chemist suggests the fuel was faulty. The teacher wonders about the pilot training and the lawyer blames it all on criminal negligence by the designer.
If you have never worked with aircraft, I wish some of you might be a little more tentative in coming to conclusions.
Last edited by Sunfish; 9th Apr 2011 at 02:11.
Join Date: Sep 2010
Location: earth
Posts: 1,341
Likes: 0
Received 0 Likes
on
0 Posts
We make estimates based on experiment as to how long a component will last, then we apply safety factors to that estimate. When the aircraft is in service, we monitor all failures and continue testing to refine those estimates based on experience. In more than one case I dealt with, we actually reduced the number of failures by extending the time in service limits for an assembly, see if you can work out why.
the stress history determines when the component will fracture
Inspection processes will improve as will manufacturing processes, at the same time manufacturers will develope products that push limits, use new alloys that will fail as they push limits. Over time we develope new inspections and servicable on wing times.