Southwest 737 lands at Yeager Airport after hole in fuselage
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As a Airframe Inspector i have to carry out crown skin inspections but have never come across ref material indicating the possibility of a localised diff pr causing the scenario you describe.
That said, Old Engineer might be onto something! But it may be more to do with varying pressures, than an absolute exterior reduction. At some fuse attitudes and Mach Nos, it's possible slight (but long term, sustained) buffeting in that area, may have contributed.
Also, without going into calcs, my gut feeling is there could be a greater pressure reduction by far from the fuse running at a small angle of attack (which it can well do early in the cruise) than from any induced flow from that fin l.e.
Something else...looking at the bonded doubler and the tear strap/strip, and the area ratios of frame to panel. How much difference in fuse mass to just make the whole thing one thickness.. if that is typical of the frame pitch and longitudinal stiffeners.Yes, 200 lb is one passenger, but by heck, what a lot of fuss for a few oz per sq.ft
Remember, the DH Comet fuse skins were about 20~22 swg, just one gauge thicker and De Havilland might still be a name in Commercial Jet Transport Aviation
That said, she's a strong'un to hang together so well after that. And corrosion doesn't look to come into the equation. Old Engineer's belling-edge stress theory sounds good to me.
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wondering which airframe builder built the toughest airframes (western world)
Tough Old Bird
My former employer proposed to lease a rather high-time DC-8 for a government contract. Based on a thorough inspection, the inspectors declared that they couldn't determine from the major structural members that the airframe had any time on it at all. Clearly a tough old bird.
P.S.: DC_ATE are you out there?
P.S.: DC_ATE are you out there?
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i dont see how boeing would use a 0.032" skin panel in this area?
thats 0.8 mm thick. if you put a 3/16 or 5/32" countersunk rivet in that
you would have a knife edge
i would expect 3.2 mm skin (maybe min of 1.6 mm).
Ive not seen bonded aluminium skin panels used before, ive worked
in this business 12 yrs as stress engineer for airbus/gulfstream
You can start of with 3.2 mm 2024 T3 or T42 (i think T42 is for stretch forming i.e. double curvature) then use chemi etching
to reduce the thickness in some areas.
Or you could use 1.6 or 1.2 mm, then use rivetted doublers of similar thicknes to increase the thickness if you didnt want to chemi etch
i.e. at window or a door cutout, there are doubler or reinforcement
patches (which is layers of skin rivetted or bolted on, to increase
the overall thickness)
thats 0.8 mm thick. if you put a 3/16 or 5/32" countersunk rivet in that
you would have a knife edge
i would expect 3.2 mm skin (maybe min of 1.6 mm).
Ive not seen bonded aluminium skin panels used before, ive worked
in this business 12 yrs as stress engineer for airbus/gulfstream
You can start of with 3.2 mm 2024 T3 or T42 (i think T42 is for stretch forming i.e. double curvature) then use chemi etching
to reduce the thickness in some areas.
Or you could use 1.6 or 1.2 mm, then use rivetted doublers of similar thicknes to increase the thickness if you didnt want to chemi etch
i.e. at window or a door cutout, there are doubler or reinforcement
patches (which is layers of skin rivetted or bolted on, to increase
the overall thickness)
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This a/c was 15 yrs old, and probably due for its 3rd D-check, which is
a major overhaul. The paint is remove and skins/splice, joints fittings
lugs, panels are inspected for cracks.
The cabin is subject to around 9 psi of pressure every flight. Producing
very large hoop tensile stresses in the fuselage skins.
The more pressurisation cycles you have, the lower the fatigue life for the a/c
Cracks in the fuselage skins are permitted provided they dont reach a
critical length between inspection intervals. A normal fatigue spectrum
takes this into account, and is used when calculating what skin thickness
to use in that area. If there are skin or lap joints then these contribute
to the fatigue life. Your more likely to have a crack start at a joint (because
of hole), than in a region where there are no holes.
My opinion is that the age/ the fatigue spectrum and maintenance schedule of this a/c could have played a part.
I.e. maybe sustained short flights (high number of pressure cycles), or a number of heavy landings (which put high longitudinal tension loads in the upper skin).
who know maybe southwest were running to close to the bone on when to
do the next D-check.
a major overhaul. The paint is remove and skins/splice, joints fittings
lugs, panels are inspected for cracks.
The cabin is subject to around 9 psi of pressure every flight. Producing
very large hoop tensile stresses in the fuselage skins.
The more pressurisation cycles you have, the lower the fatigue life for the a/c
Cracks in the fuselage skins are permitted provided they dont reach a
critical length between inspection intervals. A normal fatigue spectrum
takes this into account, and is used when calculating what skin thickness
to use in that area. If there are skin or lap joints then these contribute
to the fatigue life. Your more likely to have a crack start at a joint (because
of hole), than in a region where there are no holes.
My opinion is that the age/ the fatigue spectrum and maintenance schedule of this a/c could have played a part.
I.e. maybe sustained short flights (high number of pressure cycles), or a number of heavy landings (which put high longitudinal tension loads in the upper skin).
who know maybe southwest were running to close to the bone on when to
do the next D-check.
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there are only two wide body manufacturers left Boeing and Airbus
Boeing 777, 767, 747
Airbus: A380, a340, a330, a310, a300
lockheed tristar? i dont think there are many of them still flying pax
commerically most of them and DC10 are freighters now
Boeing 777, 767, 747
Airbus: A380, a340, a330, a310, a300
lockheed tristar? i dont think there are many of them still flying pax
commerically most of them and DC10 are freighters now
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Well, that BA 777 stood up well after dropping into Heathrow at something like12 to 15 ft/s ?
There may be the odd DC-6 still flying commerically (fire suppression in Canada?)
There may be the odd DC-6 still flying commerically (fire suppression in Canada?)
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And now for something completely different
OK we have had the usual "this brand is better" crap, and even had criticism of Airbus over a Boeing failure. Lets have an objective look at the evidence, in particular from the NTSB pictures.
Firstly, there is clearly a bonded reinforcement filling the entire bay but not splicing to the bonded fail-safe strap. This suggests that there was a known problem in the area. No manufacturer adds material as a design feature or a modification unless there is an issue. If there was not an issue discovered in later service or testing subsequent to certification testing, then the increase in section would not have have been formed in initial production, eliminating the need to add the doubler.
Secondly, the location of the crack that led to the failure is along the side of the fore-aft top fail-safe strap and also between the bonded reinforcement mentioned above. This is the thinnest section of skin between the skin+fail safe strap and skin+doubler.
Thirdly, the skin would have been designed to take the pressure loads on a pd/2t basis learnt in second year engineering. There would also be adequate design roigour (rigor for our American cousins) to allow for the ground-air-ground fatigue cycle.
Fourthly, the change in crack direction from parallel to the fore-aft structure to nearly parallel to the hoop fail-safe strap is typical of acoustic fatigue crack growth. It is not true flight-cycle fatigue or it would not change direction. The fore-aft stress is half that of the hoop stress plus or minus the fuselage bending stresses at the crown. Also, if it was true pressure cycle fatigue, it would also probably have initiated at the location of the high stress concentrations at the fasteners, not in the adjacent skin.
My assessment is that this is a case of acoustic fatigue (due to air flow disturbance at the front of the vertical stabiliser as already discussed). This was known to the OEM because of the evidence displayed by modification to incorporate the doubler. The reinforcement was added as a modification to change the local natural frequency, but resulted in a change to the local vibration mode to another mode that resulted in the small area between the doubler and the fail-safe strap being the critical fatigue area.
Acoustic fatigue is a strange beast that is not necessarily responsive to simple analysis. I have personal experience where a number of modifications changed the natural frequency to the next harmonic, or changed the displacement mode to a worse form with a consequent shortening of structural life compared to leaving it alone.
The solution is to incorporate an adhesively bonded ring-doubler on the INSIDE of the structure to eliminate the structurally weak area between the doubler and the fail-safe strap. It would also eliminate out-of plane bending associated with the change in section between the douber and the fail-safe strap. I would incorporate a damping layer as part of the modification. I would also strongly recommend an acoustic analysis of my suggestion, and an assessment of the effect on the fail-safe certification basis of the increased stiffness on adjacent bays.
Can I have STC rights for the above suggested modification? If so, my address is at my web site adhesionassociates.com
I hope this adds a level of structural discussion over and above the usual brand focussed tripe.
Regards
Blakmax
Firstly, there is clearly a bonded reinforcement filling the entire bay but not splicing to the bonded fail-safe strap. This suggests that there was a known problem in the area. No manufacturer adds material as a design feature or a modification unless there is an issue. If there was not an issue discovered in later service or testing subsequent to certification testing, then the increase in section would not have have been formed in initial production, eliminating the need to add the doubler.
Secondly, the location of the crack that led to the failure is along the side of the fore-aft top fail-safe strap and also between the bonded reinforcement mentioned above. This is the thinnest section of skin between the skin+fail safe strap and skin+doubler.
Thirdly, the skin would have been designed to take the pressure loads on a pd/2t basis learnt in second year engineering. There would also be adequate design roigour (rigor for our American cousins) to allow for the ground-air-ground fatigue cycle.
Fourthly, the change in crack direction from parallel to the fore-aft structure to nearly parallel to the hoop fail-safe strap is typical of acoustic fatigue crack growth. It is not true flight-cycle fatigue or it would not change direction. The fore-aft stress is half that of the hoop stress plus or minus the fuselage bending stresses at the crown. Also, if it was true pressure cycle fatigue, it would also probably have initiated at the location of the high stress concentrations at the fasteners, not in the adjacent skin.
My assessment is that this is a case of acoustic fatigue (due to air flow disturbance at the front of the vertical stabiliser as already discussed). This was known to the OEM because of the evidence displayed by modification to incorporate the doubler. The reinforcement was added as a modification to change the local natural frequency, but resulted in a change to the local vibration mode to another mode that resulted in the small area between the doubler and the fail-safe strap being the critical fatigue area.
Acoustic fatigue is a strange beast that is not necessarily responsive to simple analysis. I have personal experience where a number of modifications changed the natural frequency to the next harmonic, or changed the displacement mode to a worse form with a consequent shortening of structural life compared to leaving it alone.
The solution is to incorporate an adhesively bonded ring-doubler on the INSIDE of the structure to eliminate the structurally weak area between the doubler and the fail-safe strap. It would also eliminate out-of plane bending associated with the change in section between the douber and the fail-safe strap. I would incorporate a damping layer as part of the modification. I would also strongly recommend an acoustic analysis of my suggestion, and an assessment of the effect on the fail-safe certification basis of the increased stiffness on adjacent bays.
Can I have STC rights for the above suggested modification? If so, my address is at my web site adhesionassociates.com
I hope this adds a level of structural discussion over and above the usual brand focussed tripe.
Regards
Blakmax
Acoustic Fatigue
blakmax
agree with you that's what it looks like.
It's natures way of telling you that the metal is excess structure
I've seen it addressed simply by removing the piece of metal that cracks in the first place. Unfortunately this has its drawbacks when you're trying to keep air inside
agree with you that's what it looks like.
It's natures way of telling you that the metal is excess structure
I've seen it addressed simply by removing the piece of metal that cracks in the first place. Unfortunately this has its drawbacks when you're trying to keep air inside
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Blakmax suggests the bonded doubler is not a repair, but an attempt to detune that panel, as if there were a already known acoustic/resonance problem there.
I seem to concurr
I seem to concurr
That said, Old Engineer might be onto something! But it may be more to do with varying pressures, than an absolute exterior reduction. At some fuse attitudes and Mach Nos, it's possible slight (but long term, sustained) buffeting in that area, may have contributed.
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there is no doubt in my mind that
Douglas built the toughest/strongest airliners. (and the best flying for the pilot)
That Lockheed was a real innovator
And that the world is worse off for these two companies to be out of the airliner building biz.
Douglas built the toughest/strongest airliners. (and the best flying for the pilot)
That Lockheed was a real innovator
And that the world is worse off for these two companies to be out of the airliner building biz.
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Use of bonded repairs
It is not common practice to use bonded aluminium panels as repairs.
The simplest repair would be to rivet an external doubler/patch
to the skin.
The simplest repair would be to rivet an external doubler/patch
to the skin.
I have over thirty seven years experience with adhesive bonded repairs to metallic military aircraft. In one example, bonded composite patches were used to repair widespread stress corrosion cracking in C-130E wing planks. These repairs enabled the RAAF to be the only operator in the world to fly the C-130E through its life of type without replacing the wing planks. That has been estimated to have saved an audited AUD130 million. The USAF used bonded composite patches to repair C-141 and that saved billions compared to replacing the wing skins. Mechanical repairs were not an option because they could not meet damage tolerance requirements and did not provide adequate restraint of crack growth.
With skin as thin as the area in question on the 737, I would bet the family jewels that I could design a bonded repair that would never fail and would be stronger than the metal itself. In comparison, in a mechanically fastened repair the joint strength will always be limited by the stress concentrations caused by the fastener. Lets be clear. If tested to failure, an appropriately designed and correctly processed adhesive bond will break in the metal outside the repair. A mechanical repair would break at the fastener line at a much lower load.
There is a significant amount of data on bonded repairs to metallic structure. I suggest you look at Adhesion Associates for a start.
Regards
blakmax
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That Lockheed was a real innovator
And that the world is worse off for these two companies to be out of the airliner building biz.
And that the world is worse off for these two companies to be out of the airliner building biz.
De Havillands were Redux bonding skins and stringers in 1950 on jet airliners
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And Hughes used multi layer Birch lumber laminates with Phenolic resin.
deHavilland. Mosquito?
Basically, lumber is a two phase matrix like CFRP, are we 'regressing'?
Rivets are older than all dirt.
deHavilland. Mosquito?
Basically, lumber is a two phase matrix like CFRP, are we 'regressing'?
Rivets are older than all dirt.
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most repairs are done by the airline approved maintenance ppl.
They only contact the a/c manufacturer when theyve got some concerns
or are unsure what to do.
Bonded composite repairs, im sure will be the norm once the a350 and 787
are well into service. Ive heard about them talking about the patches
for the 787.
I guess the conditions have to exactly right (i.e. the curing process
the mixing materials) in order that when the patch is bonded on, it
will meet the allowables that you specify it will meet.
So what Fsu can you achieve with the adhesive?
I guess the main people you should be trying to convince are the airlines
and the certification authorities, if you say it saves money? Its only a cultural thing at the end of the day.
it doesnt take long however to make a patch, drill off 15-20 holes and install
the rivets. And at least you are 100% confident that the the rivet allowable will
be spot on.
ill look at your website
cheers
They only contact the a/c manufacturer when theyve got some concerns
or are unsure what to do.
Bonded composite repairs, im sure will be the norm once the a350 and 787
are well into service. Ive heard about them talking about the patches
for the 787.
I guess the conditions have to exactly right (i.e. the curing process
the mixing materials) in order that when the patch is bonded on, it
will meet the allowables that you specify it will meet.
So what Fsu can you achieve with the adhesive?
I guess the main people you should be trying to convince are the airlines
and the certification authorities, if you say it saves money? Its only a cultural thing at the end of the day.
it doesnt take long however to make a patch, drill off 15-20 holes and install
the rivets. And at least you are 100% confident that the the rivet allowable will
be spot on.
ill look at your website
cheers
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Bonded repairs
Hi skin-splice.
Maybe we need to start a new thread because this is getting off the subject. I'll let the moderator PM me if there is interest. I would be only too happy to start the ball rolling, probably under the "Engineers and Technicians" stream.
Bonded repairs to composite structure will definitely be the predominant repair if anyone has any inteligence at all. Unfortunately my understanding is that OEMs dumb-down SRMs to enable airlines in Bogloviastan to still perform repairs so much of the repairs to new composite structures will be by mechanical fastening, despite the disasterous strength loss due to tear-out weakness of composites.
I have great concerns about some SRM repairs. Injection repairs for disbonds can never under any circumstances provide an ounce of structural integrity for example, yet they remain the standard repair in SRMs. I defy anyone to show me evidence that they do anything except fill the air gap so that they pass NDI and give the technician a warm fuzzy feeling that he has repaired the defect. I have hundreds of examples including some where this OEM repair method has resulted in in-flight failures of large structures.
You ask what Fsu can be achieved by an adhesive bond. The answer is as much as the metal can sustain up to a thickness of 0.15 in for most high strength aluminium alloys. Firstly, understand that lap shear strength is a meaningless parameter which has absoultely nothing to do with bond strength, unless your bond is between 0.062 2024-T3 aluminium and you have used a stupid overlap length of 0.5 inches. Change the material, the overlap length or the thickness of the adherends and you get a different value.
It is actually possible to calculate the load capacity of the bond (the strength of the bond if the adherend failure is ignored) using equations derived by John Hart-Smith in the early 1970's. These equations deal with dissimilar materials, different adherend thicknesses and even thermal stresses. If that load capacity is greater than the unnotched strength of the adherends, then the adherends will fail first provided the overlap length is adequate. It will be physically impossible to exceed the strength of the adhesive because the adherends will fail first.
Next is the processing. You are correct that this is important and there are three factors; the surface preparation method, the temperature measurement and control processes and contamination control. Contamination is well known as a source of bond failure and simple methods can significantly reduce the risks. Temperature measurement and control for hot-bonding is currently a disaster waiting to happen. Almost every SRM directs the use of a single heater blanket irrespective of heat sinks and heat loss paths. They also specify a set array of thermocouples (four thermocouples arranged at 90 degrees apart). These directions result in either overheating of the structure or undercure of the adhesive and subsequent interfacial failure. I am aware of a 767 in the Pacific with a trailing edge flap corner repair performed by an authorised repair station that has fallen off the aircraft at least four times. My bet is that the adhesive has been heated too slowly causing the polymers to crosslink before the adhesive has flowed sufficiently to wet the surface. The mechanism is the same as using adhesive which is out of life. The failure is at the interface.
By far the most critical is the surface preparation method. Here is the real problem. Do you know that current certification methods for aircraft structures requires demonstration of static strength and fatigue resistance, but current regulations DO NOT require demonstration of long term bond durability. The only requirement is that bonding processes must produce a "sound" structure. What is a sound structure? If it passes NDI is that sound? If it passes static strength and fatigue testing is that sound? The anwer is a definite NO!
I could make a bonded structure which passes static strength and fatigiue tests in the short term, but will fail in long term service. The mechanism invoilved is that the chemical bonds formed at the time the adhesive is cured are sufficient to enable static strength and fatigue performance. However, in service these bonds degrade usually by hydration of the metal oxides. To enable the oxide surface to hydrate the chemical bonds at the interface dissociate leading to interfacial disbonding. The secret to a successful bond is to treat the surface at the time of bonding with chemicals which prevent hydration.
Now because the FARs, JARs etc. do not mandate testing to demonstrate resistance to hydration, we continue to see interfacial failures and loads well below even limit load, let alone ultimate load. Even worse, the bonding methods contained in SRMs are among the worst offenders. The only processes I would trust are phosphoric acid anodising (below 85F or 29C) the RAAF grit blast and silane process and Boeing's sol-gel process. Pasajel, alodine and many other proprietary processes are useless.
All of this is contained in the FAA publication DOT/FAA/AR – TN06/07, Apr 2007 available throught eh FAA Tech Center. Unfortunately because of homeland security, people outside the US (inluding myself as author) can not access the docoment without a request by email. I can send a draft copy to anyone on request.
Regards
blakmax.
Now MOD how about a new thread on adhesive bonding technology?
Maybe we need to start a new thread because this is getting off the subject. I'll let the moderator PM me if there is interest. I would be only too happy to start the ball rolling, probably under the "Engineers and Technicians" stream.
Bonded repairs to composite structure will definitely be the predominant repair if anyone has any inteligence at all. Unfortunately my understanding is that OEMs dumb-down SRMs to enable airlines in Bogloviastan to still perform repairs so much of the repairs to new composite structures will be by mechanical fastening, despite the disasterous strength loss due to tear-out weakness of composites.
I have great concerns about some SRM repairs. Injection repairs for disbonds can never under any circumstances provide an ounce of structural integrity for example, yet they remain the standard repair in SRMs. I defy anyone to show me evidence that they do anything except fill the air gap so that they pass NDI and give the technician a warm fuzzy feeling that he has repaired the defect. I have hundreds of examples including some where this OEM repair method has resulted in in-flight failures of large structures.
You ask what Fsu can be achieved by an adhesive bond. The answer is as much as the metal can sustain up to a thickness of 0.15 in for most high strength aluminium alloys. Firstly, understand that lap shear strength is a meaningless parameter which has absoultely nothing to do with bond strength, unless your bond is between 0.062 2024-T3 aluminium and you have used a stupid overlap length of 0.5 inches. Change the material, the overlap length or the thickness of the adherends and you get a different value.
It is actually possible to calculate the load capacity of the bond (the strength of the bond if the adherend failure is ignored) using equations derived by John Hart-Smith in the early 1970's. These equations deal with dissimilar materials, different adherend thicknesses and even thermal stresses. If that load capacity is greater than the unnotched strength of the adherends, then the adherends will fail first provided the overlap length is adequate. It will be physically impossible to exceed the strength of the adhesive because the adherends will fail first.
Next is the processing. You are correct that this is important and there are three factors; the surface preparation method, the temperature measurement and control processes and contamination control. Contamination is well known as a source of bond failure and simple methods can significantly reduce the risks. Temperature measurement and control for hot-bonding is currently a disaster waiting to happen. Almost every SRM directs the use of a single heater blanket irrespective of heat sinks and heat loss paths. They also specify a set array of thermocouples (four thermocouples arranged at 90 degrees apart). These directions result in either overheating of the structure or undercure of the adhesive and subsequent interfacial failure. I am aware of a 767 in the Pacific with a trailing edge flap corner repair performed by an authorised repair station that has fallen off the aircraft at least four times. My bet is that the adhesive has been heated too slowly causing the polymers to crosslink before the adhesive has flowed sufficiently to wet the surface. The mechanism is the same as using adhesive which is out of life. The failure is at the interface.
By far the most critical is the surface preparation method. Here is the real problem. Do you know that current certification methods for aircraft structures requires demonstration of static strength and fatigue resistance, but current regulations DO NOT require demonstration of long term bond durability. The only requirement is that bonding processes must produce a "sound" structure. What is a sound structure? If it passes NDI is that sound? If it passes static strength and fatigue testing is that sound? The anwer is a definite NO!
I could make a bonded structure which passes static strength and fatigiue tests in the short term, but will fail in long term service. The mechanism invoilved is that the chemical bonds formed at the time the adhesive is cured are sufficient to enable static strength and fatigue performance. However, in service these bonds degrade usually by hydration of the metal oxides. To enable the oxide surface to hydrate the chemical bonds at the interface dissociate leading to interfacial disbonding. The secret to a successful bond is to treat the surface at the time of bonding with chemicals which prevent hydration.
Now because the FARs, JARs etc. do not mandate testing to demonstrate resistance to hydration, we continue to see interfacial failures and loads well below even limit load, let alone ultimate load. Even worse, the bonding methods contained in SRMs are among the worst offenders. The only processes I would trust are phosphoric acid anodising (below 85F or 29C) the RAAF grit blast and silane process and Boeing's sol-gel process. Pasajel, alodine and many other proprietary processes are useless.
All of this is contained in the FAA publication DOT/FAA/AR – TN06/07, Apr 2007 available throught eh FAA Tech Center. Unfortunately because of homeland security, people outside the US (inluding myself as author) can not access the docoment without a request by email. I can send a draft copy to anyone on request.
Regards
blakmax.
Now MOD how about a new thread on adhesive bonding technology?
Last edited by blakmax; 21st Jul 2009 at 10:43. Reason: clarification (added missing words "including some" in third para)