737NGs have cracked 'pickle forks' after finding several in the jets.
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Most of the discussion on this thread about winglets is the additional stresses that they impose on a wing. I don't think anyone has stated here that winglets are the sole possible reason for the cracking of the pickle fork There are other possible reasons why the pickle fork could crack including High CYCLES vs Hours.Folks on this thread have brought up many possibilities for the cracking of the pickle fork including improper manufacturing or assembly processes.
Any chance that repeated overweight or hard landings could have been the cause
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737 ng picklefork cracks actual cause
The actual cause of the pickle fork cracks is because of the the holes being over drilled by around 6 thousands. The holes are grossly oversize and the bolts that run through the part and the fail safe strap are not supporting the part. The bolts are supposed to be snug fit so when a plane lands the bolts and part share equal stress and not have a stress issue. The bolts are not touching the sides of the holes hardly at all therefore any gap between the bolt and the part contributes to all stress being put on the fork part and fail safe strap, which has holes where the bolts are.
The actual cause of the pickle fork cracks is because of the the holes being over drilled by around 6 thousands. The holes are grossly oversize and the bolts that run through the part and the fail safe strap are not supporting the part. The bolts are supposed to be snug fit so when a plane lands the bolts and part share equal stress and not have a stress issue. The bolts are not touching the sides of the holes hardly at all therefore any gap between the bolt and the part contributes to all stress being put on the fork part and fail safe strap, which has holes where the bolts are.
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IF- BIG IF the claim that holes are-were oversize as claimed ( some worker- bee who is involved in repair-replace would know ) then ther problem is almost exactly as I postulated back on sept 11 - see below. depending on the defined process - which SHOULD have required the holes on final assembly in that area be properly drilled- reamed- inspected as a minimum ( and skipped or pencil whipped ) - or IMHO been properly coldworked, then the problem could be fleetwide. Although it would be awkward - it **may** be possible to schedule lower cycle planes to have those holes in that area reamed, coldworked, and proper interference fit fasteners installed before say 15,000 to 20,000 cycles.
So much for faster - cheaper- - by god we didn't do that on the Boeing- westervelt B-1, and its just an extra time consuming thing to do on assembly and meet or beat the BAR ( chart ). Just a SWAG- but I'll bet that process IF done- specified up thru about 1998-99 may have been dropped.
So much for faster - cheaper- - by god we didn't do that on the Boeing- westervelt B-1, and its just an extra time consuming thing to do on assembly and meet or beat the BAR ( chart ). Just a SWAG- but I'll bet that process IF done- specified up thru about 1998-99 may have been dropped.
It will be interesting to see or find out just where the cracks are. Obviously at this time, its just a guess. However, as a GUESS and based on experience dealing with fastener issues in large parts on 707 and 767 ( 707 that had been in service for a long time ) and tooling for 767-here is my SWAG based on the relatively low key descriptions.
A) The cracks are probably around or spreading from Fastener holes, probably those drilled " by hand" during the LEAN manufacturing process which are less than about 3/8 in diameter.
B) As such there is of course an argument that the crack would simply progress to the next hole in the pattern ( since drilling a small hole at the ' end ' of such crack is considered to be a ' crack stopper ' - which is true for a lot of ' sheet metal ' issues.
C) again , just a guess, but for 40 plus years, thre has been available and used a three to four step process to prevent such cracks, which can be done for all sizes of hole, even large holes during fabrication while still in large tooling- drill plates , etc.
D) in general- the steps are 1) drill a hole slightly smaller than final size 2) insert a thin sleeve into hole 3) insert a special mandrel such that when pulled back thru the hole it expands the sleeve and hole. 4) Ream the hole which will usually be slightly out of round to final size.
On assembly, insert bolt as a tight fit.
E) in some cases and sizes, the same process can be used without a sleeve but with a expanding mandrel.
F) This leaves a major prestress around the hole and provides a significant improvement in fatigue life.
The process was patented by Boeing in the late 1960, and a local firm called Fatigue technology was founded- developed from the previous firm called Industrial Wire and metal forming as I recall. And major first use ( from memory ) was on AWACS.
The process- tooling has been the subject of several related patents, and is still used by virtually all aircraft manufacturers
It can be used to prevent or stop cracks from further progress.
Again MY SWAG is that to save time some $$$ - or due to a temporary lack of sleeves or just plain skipping the sequence ( hard to detect when inspection is only on final hole size )
So depending on location and accessibility, the fix would be to remove bolt, expand hole, ream hole, insert new oversize bolt and voila, a terminating fix.
Just have to wait and see- If someone has access to documentation as to real issue and location, would be interesting to see how close I came
A) The cracks are probably around or spreading from Fastener holes, probably those drilled " by hand" during the LEAN manufacturing process which are less than about 3/8 in diameter.
B) As such there is of course an argument that the crack would simply progress to the next hole in the pattern ( since drilling a small hole at the ' end ' of such crack is considered to be a ' crack stopper ' - which is true for a lot of ' sheet metal ' issues.
C) again , just a guess, but for 40 plus years, thre has been available and used a three to four step process to prevent such cracks, which can be done for all sizes of hole, even large holes during fabrication while still in large tooling- drill plates , etc.
D) in general- the steps are 1) drill a hole slightly smaller than final size 2) insert a thin sleeve into hole 3) insert a special mandrel such that when pulled back thru the hole it expands the sleeve and hole. 4) Ream the hole which will usually be slightly out of round to final size.
On assembly, insert bolt as a tight fit.
E) in some cases and sizes, the same process can be used without a sleeve but with a expanding mandrel.
F) This leaves a major prestress around the hole and provides a significant improvement in fatigue life.
The process was patented by Boeing in the late 1960, and a local firm called Fatigue technology was founded- developed from the previous firm called Industrial Wire and metal forming as I recall. And major first use ( from memory ) was on AWACS.
The process- tooling has been the subject of several related patents, and is still used by virtually all aircraft manufacturers
It can be used to prevent or stop cracks from further progress.
Again MY SWAG is that to save time some $$$ - or due to a temporary lack of sleeves or just plain skipping the sequence ( hard to detect when inspection is only on final hole size )
So depending on location and accessibility, the fix would be to remove bolt, expand hole, ream hole, insert new oversize bolt and voila, a terminating fix.
Just have to wait and see- If someone has access to documentation as to real issue and location, would be interesting to see how close I came
In a clamped joint with high tensile bolts, the bolts are not (supposed to be) subjected to shear loads, which are carried by friction between the surfaces which are clamped together by the bolts. Holes have clearance and individual bolts are not subjected to load in sequence, failing before the next bolt takes load.
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There is a chance of calculating a bolt joint where - as Nonsense said- bolts do not carry shear. With shear carrying bolts and variable hole diameters and varying location tolerances not a chance. In spite of the calculation capacity existing experimental data is still used, that database is huge.
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In a clamped joint with high tensile bolts, the bolts are not (supposed to be) subjected to shear loads, which are carried by friction between the surfaces which are clamped together by the bolts. Holes have clearance and individual bolts are not subjected to load in sequence, failing before the next bolt takes load.
Maybe the reason why those cracks are taken so seriously : the joints have slipped.
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I am the source lol
the source comes from what Boeing and spirit told our facility during their visits this week. Though the company I work for is not the cause of the problem forms, we took over the pickleforks parts in 2018. Spirit and Boeing have visited and seen our entire procedure from start to finish. They are very happy with our production of the pickleforks for the max 737 and the replacement forks we are building for the 737ng. We are the sole company in charge of making the forks now and the info I have about the grossly oversized holes is straight from Boeing and spirit themselves.
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In a clamped joint with high tensile bolts, the bolts are not (supposed to be) subjected to shear loads, which are carried by friction between the surfaces which are clamped together by the bolts.
An important factor in bonding, beyond the capacity of the bonded joint itself, is that the bonding excludes contaminates from the joint. If the primary structural lap joint were to be depending upon friction for load carrying, what would happen when oil seeped into it? The friction would go away! You'd sure hope then that the fasteners would carry the load in shear! This can be better understood in real life, in that friction is a factor when properly torquing nuts and bolts, to assure that the designed torque is not exceeded.
In most cases, torquing specifies "dry threads", because the many different types of lubricants, if applied to the threads of the fastener being torqued, could dramatically change the achieved torque, and thus tension applied to the bolt while torquing. When experimenting, I have managed to snap off 125ksi aircraft bolts in their normal torquing range, by applying really good lubricants to the threads. Friction is very hard to predict and calculate, but the lack of friction is really easy to figure out, and quite achievable with a good lubricant. My structural designs rely on standard methodology fastener shear allowables calculated to carry the entire load of a structural lap joint in shear.
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It is a bit more complicated. In practice the load is carried (almost all) by friction, it is calculated to be ably to carry the load by bolt shear too.
If it were bolt shear only the tightening torques would be quite low, no need for high ones.
With pre drilled holes the load carrying capacity would be a lot lower than the theoretical one with in place reamed holes.
I remember one composite-metal connection where it was dimensioned for both bonded and bolts. In practice the load is/was carried by the adhesive.
If it were bolt shear only the tightening torques would be quite low, no need for high ones.
With pre drilled holes the load carrying capacity would be a lot lower than the theoretical one with in place reamed holes.
I remember one composite-metal connection where it was dimensioned for both bonded and bolts. In practice the load is/was carried by the adhesive.
In a clamped joint with high tensile bolts, the bolts are not (supposed to be) subjected to shear loads, which are carried by friction between the surfaces which are clamped together by the bolts. Holes have clearance and individual bolts are not subjected to load in sequence, failing before the next bolt takes load.
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One of the earliest versions of a tight hole filled fastener is the ' taperlok'- used extensively/ on the B-52. Then came close fit lockbolts, hi shear and similar. Then came things like rivbolt- interference fit - 'cold expansion' ( AKA coldwork ) in the 60s. Rivets were designed to expand in then hole via both squeeze and squeeze vibrate- and for some critical areas ONE SHOT installation and hole filling was controlled by die shape. ( I'm talking aeospace - since until the 60's and 70s, hot riveting was used on buildings and bridges ) and hot riveting was tried - tested ( electrical heating ) for aerospace but had other problems. One shot riveting is now done by ' electro- magnetic riveting ' and the major firm is known as electro-impact. Turns out that properly done, one shot riveting does have significant fatigue improvement.
ALL of which to say is the major- most common installation of fasteners in aerospace is to produce close or interference fit to ensure fasteners are uniformly loaded in SHEAR.
stepping down from soapbox
BTW- electro-magnetic- riveting ( one shot ) was developed and patented by Boeing in the 70's- actually developed and tested in the late 60's and used on early 747's. Electro impact came along later ( long story ) with a low voltage version and thecompany was founded on that modification.
In the late 60's, coldworking was used as a field ' repair ' on some fastener holes in high strength steel on lufthansa 707- by sending an AOG crew there with a few driils, reamers, and sleeves and mandrel. I turned down that particular trip ..
Last edited by Grebe; 16th Oct 2019 at 18:54. Reason: added a bit of background
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IF- BIG IF the claim that holes are-were oversize as claimed ( some worker- bee who is involved in repair-replace would know ) then ther problem is almost exactly as I postulated back on sept 11 - see below. depending on the defined process - which SHOULD have required the holes on final assembly in that area be properly drilled- reamed- inspected as a minimum ( and skipped or pencil whipped ) - or IMHO been properly coldworked, then the problem could be fleetwide. Although it would be awkward - it **may** be possible to schedule lower cycle planes to have those holes in that area reamed, coldworked, and proper interference fit fasteners installed before say 15,000 to 20,000 cycles.
So much for faster - cheaper- - by god we didn't do that on the Boeing- westervelt B-1, and its just an extra time consuming thing to do on assembly and meet or beat the BAR ( chart ). Just a SWAG- but I'll bet that process IF done- specified up thru about 1998-99 may have been dropped.
So much for faster - cheaper- - by god we didn't do that on the Boeing- westervelt B-1, and its just an extra time consuming thing to do on assembly and meet or beat the BAR ( chart ). Just a SWAG- but I'll bet that process IF done- specified up thru about 1998-99 may have been dropped.
As I said before the ng pickleforks were built prior to 2018. By a company. The company I work for took over the pickleforks process and only built 200 of the ng style pickleforks. The forks we built are not problematic because we never had oversized holes. We then changed over to the pickleforks for the max 737. There are no issues fleetwide because when we took over we have had numerous inspections and have never had an oversized hole. You are correct that somewhere someone was rushing and not doing proper inspections of the parts, however this was in the beginning by a smaller company that didn’t have the engineering and quality we currently have on the max pickleforks. We’re fixing their mess ups.
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Plate rivet joints are a different design case. And expanding rivets were/are used for the reason mentioned.
Steel/steel joints are mostly designed both for friction and bolt shear. Friction caries the normal loads but if something goes wrong in the chain of events the bolts still have the capacity.
In the pickle fork case at least the fork is Al, don't know about the center wing box. I have never designed a steel/Al joint but might be that the three fold difference in elastic moduli could cause some trouble. Al's low bearing strength would be a problem with that large steel bolts. I still think friction is included.
Steel/steel joints are mostly designed both for friction and bolt shear. Friction caries the normal loads but if something goes wrong in the chain of events the bolts still have the capacity.
In the pickle fork case at least the fork is Al, don't know about the center wing box. I have never designed a steel/Al joint but might be that the three fold difference in elastic moduli could cause some trouble. Al's low bearing strength would be a problem with that large steel bolts. I still think friction is included.