Bolt holes, burr and fatigue life effects

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wktaylor

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OH YEAH... 'WM'... even though I know You are kidding...

RE: "Only half the rivets need to hold, so that cuts down workload."

Question… Which half-of-the-rivets in a row should 'hold'... front half, back half, middle half, every-other, groups of '3s'... etc?

Recently my organization encountered a ridiculous problem... we discovered that one/more replacement critical upper skin panels LH & RH wings on our many Depot maintenance jets... since 2012... were installed using highly non-standard/un-documented methods to speed-up production. A high majority of the rivets were irregularly bucked... every discrepancy imaginable.

Us engineers were asked: what is safety margin if 1/2 of the rivets were installed unsatisfactorily [other than fuel leaking]. Since riveting has become so standardized over the decades... and the 'acceptable limits' were tested and validated long-ago... and since rivets are accessible for detailed visual inspection... there was never any issue [reports, standards, etc] regarding wide-scale substandard workmanship... other-than rules such as "1-in10 questionable rivet installs... randomly scattered and no [3] clustered together".

The only answer we all had was this is a critical airworthiness issue.... going back 8-years worth of Depot maintenance. Mandatory removal and replacement of millions of discrepant rivets, under urgent circumstances, will cost millions of $$s, is warranted. EVEN the 'good' rivets have to be replaced since they have been disproportionately stressed... absorbing load/strain that 'bad rivets' should have been carrying.
 
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TuweetyBird

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Google Navair NA 01-1A-1....just happened to remember this guy as I was reading through here, plenty of PDFs available...be safe.
 

wktaylor

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USAF => 1-1A-1 Engineering Handbook Series For Aircraft Repair - General Manual For Structural Repair [co-numbered 01-1A-1] Technical Orders

USAF => 1-1A-8 Engineering Manual Series - Aircraft and Missile Repair - Structural Hardware [co-numbered 01-1A-8]
 

wktaylor

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BTW... TuweetyBird for grins...

In the 1980s [~4-years], I was USAF lead structural/mechanical engineer [SA-ALC, Kelly AFB] for the T-37 Tweetybird and the A/OA-37B Dragonfly. Great experiences... lots of mishap training, though...
 

SamP

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Thanks for everyone's input. My brain is slowly "grokking." Just trying to build the right model in my brain so all the factoids I pick up fit together.

Interestingly, I learned as I read through materials that while differences in clearance may affect how much load a particular bolt takes, oversized holes don't affect ultimate failure or maximum load, but does affect fatigue failure. I also learned that maximum load drops dramatically if your don't drill perpendicular to the surface.

From my understanding, bolt joints are designed to hold pieces together, and that friction between the two pieces is the predominant factor, not shear strength of a bolt. That friction is created by the bolt preload (how hard you are cranking down on the nut). I also know that the preload on bolts are pretty low for aircraft applications, implying that it is the bolt getting loaded in shear that is doing much of the work. Is there a contradiction there?

Edit: I just read that some shear joints are designed for the bolt itself to take the shear, and not the friction between the plates. It would seem that the aircraft joints are designed according to this So maybe no contradiction.

There is also what entails joint failure. Is it the bolt shearing (then oversizing shouldn't affect the failure since it's the same bolt), or maybe slippage / deformation. There is rabbit hole I'm definitely going down...
 
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Matt G.

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No contradiction. In reality, the friction does do some of the work, but in analysis of aircraft structure it is always neglected (perhaps I should say almost always...someone here probably has an exception I have yet to encounter in my career), and assumed 100% of the load is transmitted between the two halves of the joint through fastener shear.

The failure mode of the bolted joint will depend on a lot of things, such as the materials being joined together, their thicknesses, the type, material, and diameter of the fastener, the thickness of the stackup of parts in the joint, and fastener spacing and distance from a free edge, among other things. Oversizing will increase the bearing strength of the sheet at that location, as well as the fastener shear, but having one larger fastener can cause issues in a fatigue critical area. If it is significantly oversized and the fasteners are close together, net area across the joint may become an issue, depending on the structure and loading.

The nice thing is that if you stick with recommended design practices and choose fasteners and materials that are commonly used for aircraft, there are joint allowables that have been developed by test that have many of these things built in.
 

robertbrown

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Doing mod & repair work on large transport aircraft, a lot of poor workmanship would be uncovered as structure was disassembled. Someone remarked that instead of analyzing for a "rogue flaw" that was a .010" long crack in a fastener hole, we should assume that the rogue flaw was a large chisel mark in the hole. I would say that workmanship on the wing was better; critical holes there were classified as "Fracture & Fatigue Critical" and had more stringent inspections.
 

gtae07

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This mechanical engineer sees a bunch of individual issues that are taught to mechanical engineers:
I wish those things were taught to aerospace engineers, too. Unfortunately it seems the AE schools (or at least mine) see fit to leave that kind of "practical stuff" for employers/industry to teach. Gets in the way of theory and research, you see...


Reading accident reports for cases caused by minor manufacturing defects is sobering...
 

robertbrown

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I wish those things were taught to aerospace engineers, too. Unfortunately it seems the AE schools (or at least mine) see fit to leave that kind of "practical stuff" for employers/industry to teach. Gets in the way of theory and research, you see...


Reading accident reports for cases caused by minor manufacturing defects is sobering...
I had a boss who had been an AE student when he got a job as a draftsman. He couldn't manage working and going to school at the same time so he dropped out. At that time, they didn't have any problem with promoting draftsmen to design engineers and he finished up his career as an engineering supervisor. He said that it had worked out well because most of the AE's wound up in preliminary design or aerodynamics while more ME's got to work with the aircraft structure and he preferred being close to the aircraft. They don't tell you all that in school. We had several engineering coop students who changed majors after working with us but I'd rather blame that on pushing STEM on everyone regardless of individual personalities than on us being like Dilbert's office.
 

Unnikrishnan S

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Manual riveting always lead to decrease in fatigue life. You can do cold working to compensate the reduction in fatigue life.
 

wktaylor

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SamP...

‘Definition of failure’... That is a subject of fairly deep discussion. ‘Testing is our friend’... as well as established industry, DoD and corporate standards.

MG...

Your last post... some good points noted... and good design advice!

U.S. ...

Manual riveting always lead to decrease in fatigue life. You can do cold working to compensate the reduction in fatigue life.”

NOT sure Your experience... but mine is completely different... especially in the range/size of GA aircraft being discussed here... But I would tend to included attack/fighter and tanker/transport/bomber in this statement [with exceptions of course].

Disagree... been in the business since 1979... working FTI-SS-CX [FTI Split-Sleeve hole cold expansion] since 1984 for USAF fatigue-life enhancement of wing attach lugs.

Solid riveting with aluminium fasteners up to 1/4-D [6.35 mm] Dia nominal or OS, is not just practical it is doable indefinitely with good overall fatigue results... AS LONG AS well-established good practices for hole-drilling, deburring, clamp-up, set-up, and tools/tooling are used.

NOTE1. Benefit of automated assembly is the mind-boggling consistency over millions of fasteners day-in-day-out. Hand assembly suffers from ‘human factors’ mental/physical . With advent of Ti-Cb, Monel, A286 or special hybrid[multi-alloy] Ti shear-pins... there is the advantage of higher strengths with the interference developed by bucking or squeezing.

I have never seen hole-cold-working used for bolts/lockbolts/Hi-Loks/Blind bolts in the diameter ranges and applications for solid rivets... usually pin-pull lockbolts, Hi-Loks/Hi-Tigues [or newer variations of these type] in light-medium-high [depends] interference-fit installs.

In the larger Dia fastener installs where bucking is impractical and high load transfer is required, the FTI-SS-CX is used sparingly as costs go skyward fast... and usually when testing and analysis indicates a fatigue cracking potential ‘hot-spot’... rarely for a high % of fasteners [although very special cases abound]. In these reamed holes we normally install appropriate bolts/lockbolts/Hi-Loks/blind-bolts. cost is high for this value-added process.

IN EVERY CASE... assembly workmanship... especially drilling/reaming/deburring/clamp-up are huge factors in joint strength durability. This is one of the primary reasons mass produced WWII aircraft are not around... quick/dirty/rough assembly for an expected short/rough life... then throw-away.

War-story...
Years ago I reviewed an industry tech report related to the tear-down inspection of [2] very high time [+/-120,000-Hr] identical model 727s. In both cases, an amazing observation was noted... in the mostly riveted upper cabin/crown structure [LH, RH, straight sections]. One side had a significant number of large defects... cracks and associated corrosion... while the mirror image side had minimal numbers of much smaller cracks/corrosion defects... a truly astounding and mind-boggling contradiction. Why such stark differences side-to-side? After months of review it became evident that the side with ‘least damage’ had an appearance of ‘better quality of workmanship in details’... holes, deburring, cleaning/chip removal, part fit-up, etc. The mystery deepened.

Then a researcher got a bright idea... interview the shop mechanics who built these sections to try and identify variables from their perspective. It was immediately evident that the ‘same’ crews had worked together day-in/day-out for over a decade doing exactly the ‘same’ fuselage sections sides [LH or RH] under ‘different’ team leads and managers... the reason was finally apparent! One team emphasized ‘by-the-book/good-enough’ production workmanship with old-school leads and managers who’d mostly worked together since WWII and had a 'gitter-done and take a smoke’ philosophy... while the other-team emphasized ‘by-the-book/better-than/methodical ' production workmanship with a noticeably different team-pride over their attention to ‘details’ with younger leads and managers who had no WWII assembly experience. This study had multiple benefits/outcomes. The fatigue/corrosion benefit of ‘attention to details’ was significant. Humans tend to perform poorly/erratically over many years. Leadership and willingness that drives the work to a higher standard cannot be overstated. Automated Assy has become the ‘norm’ to eliminate/minimize human factors.
 

wktaylor

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NOTE. I have Bede's original book on building the BD-4... talks about riveting and the effects on airframe service-life. He emphasized the need for deburring and primer by giving the WWII production example and results of current industry testing. There was an almost 10-fold increase in trouble-free service-life by doing these [2] simple practices.

Deburring and applying primer to sheet metal was rare during WWII mass production due to the time/$ involved. Throw-away airplanes were being built at an unbelievable pace for war. It showed. Most surplus'ed war-birds had to be torn down to the-bone for attention to details omitted at the factory... such as deburring and primer... otherwise they became 'real-junk' in ~10-years.
 
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robertbrown

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SamP...

‘Definition of failure’... That is a subject of fairly deep discussion. ‘Testing is our friend’... as well as established industry, DoD and corporate standards.

MG...

Your last post... some good points noted... and good design advice!

U.S. ...

Manual riveting always lead to decrease in fatigue life. You can do cold working to compensate the reduction in fatigue life.”

NOT sure Your experience... but mine is completely different... especially in the range/size of GA aircraft being discussed here... But I would tend to included attack/fighter and tanker/transport/bomber in this statement [with exceptions of course].

Disagree... been in the business since 1979... working FTI-SS-CX [FTI Split-Sleeve hole cold expansion] since 1984 for USAF fatigue-life enhancement of wing attach lugs.

Solid riveting with aluminium fasteners up to 1/4-D [6.35 mm] Dia nominal or OS, is not just practical it is doable indefinitely with good overall fatigue results... AS LONG AS well-established good practices for hole-drilling, deburring, clamp-up, set-up, and tools/tooling are used.

NOTE1. Benefit of automated assembly is the mind-boggling consistency over millions of fasteners day-in-day-out. Hand assembly suffers from ‘human factors’ mental/physical . With advent of Ti-Cb, Monel, A286 or special hybrid[multi-alloy] Ti shear-pins... there is the advantage of higher strengths with the interference developed by bucking or squeezing.

I have never seen hole-cold-working used for bolts/lockbolts/Hi-Loks/Blind bolts in the diameter ranges and applications for solid rivets... usually pin-pull lockbolts, Hi-Loks/Hi-Tigues [or newer variations of these type] in light-medium-high [depends] interference-fit installs.

In the larger Dia fastener installs where bucking is impractical and high load transfer is required, the FTI-SS-CX is used sparingly as costs go skyward fast... and usually when testing and analysis indicates a fatigue cracking potential ‘hot-spot’... rarely for a high % of fasteners [although very special cases abound]. In these reamed holes we normally install appropriate bolts/lockbolts/Hi-Loks/blind-bolts. cost is high for this value-added process.

IN EVERY CASE... assembly workmanship... especially drilling/reaming/deburring/clamp-up are huge factors in joint strength durability. This is one of the primary reasons mass produced WWII aircraft are not around... quick/dirty/rough assembly for an expected short/rough life... then throw-away.

War-story...
Years ago I reviewed an industry tech report related to the tear-down inspection of [2] very high time [+/-120,000-Hr] identical model 727s. In both cases, an amazing observation was noted... in the mostly riveted upper cabin/crown structure [LH, RH, straight sections]. One side had a significant number of large defects... cracks and associated corrosion... while the mirror image side had minimal numbers of much smaller cracks/corrosion defects... a truly astounding and mind-boggling contradiction. Why such stark differences side-to-side? After months of review it became evident that the side with ‘least damage’ had an appearance of ‘better quality of workmanship in details’... holes, deburring, cleaning/chip removal, part fit-up, etc. The mystery deepened.

Then a researcher got a bright idea... interview the shop mechanics who built these sections to try and identify variables from their perspective. It was immediately evident that the ‘same’ crews had worked together day-in/day-out for over a decade doing exactly the ‘same’ fuselage sections sides [LH or RH] under ‘different’ team leads and managers... the reason was finally apparent! One team emphasized ‘by-the-book/good-enough’ production workmanship with old-school leads and managers who’d mostly worked together since WWII and had a 'gitter-done and take a smoke’ philosophy... while the other-team emphasized ‘by-the-book/better-than/methodical ' production workmanship with a noticeably different team-pride over their attention to ‘details’ with younger leads and managers who had no WWII assembly experience. This study had multiple benefits/outcomes. The fatigue/corrosion benefit of ‘attention to details’ was significant. Humans tend to perform poorly/erratically over many years. Leadership and willingness that drives the work to a higher standard cannot be overstated. Automated Assy has become the ‘norm’ to eliminate/minimize human factors.
I've seen one study that found stress corrosion after holes in 7075-T6 were cold worked. Not a problem with over-aged tempers but better be careful with the old standbys like 2024-T3/-T4. The WWII production philosophy got passed on to a lot of guys who came into the business during the 1950's-60's. One manager said he'd drive the tug over anyone who tried to stop an airplane from moving on schedule. At least those old timers knew what they were doing and when they were taking shortcuts. There's a new generation of mangers who don't know which end of the airplane to point down the runway but they know budget and schedule.
 

wktaylor

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RobertBrown… FTI-SS-CX has high value for retarding fatigue-crack initiation and growth and retarding stress corrosion cracking [SCC] in thick section holes... 2xxx-T3x and 7xxx-T6x alloys... but the finish-reamed-holes need careful deburring and the holes/surrounding-surfaces need high quality corrosion protective coatings [protection by isolation]. Better-yet, surrounding surfaces benefit tremendously from [glass/ceramic-bead or steel] shot-peening... before finishing. Used all these 'methods' for extending life of old structure.

See attached 'shop awareness briefing I obtained +30-years ago... lots of valuable info with cartoons. The essence of this hand-out was extracted from a detailed DoD document. Dammm. Older documents mixed photos, illustrations, cartoons, tables, charts into the text... to break-up the learning into 'short bites'. I really miss useful cartoons and hand-sketch illustrations not present in current-generation documents... they provide insight with humor. These editorial characteristics help readers retain the essence of the information!
 

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robertbrown

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RobertBrown… FTI-SS-CX has high value for retarding fatigue-crack initiation and growth and retarding stress corrosion cracking [SCC] in thick section holes... 2xxx-T3x and 7xxx-T6x alloys... but the finish-reamed-holes need careful deburring and the holes/surrounding-surfaces need high quality corrosion protective coatings [protection by isolation]. Better-yet, surrounding surfaces benefit tremendously from [glass/ceramic-bead or steel] shot-peening... before finishing. Used all these 'methods' for extending life of old structure.

See attached 'shop awareness briefing I obtained +30-years ago... lots of valuable info with cartoons. The essence of this hand-out was extracted from a detailed DoD document. Dammm. Older documents mixed photos, illustrations, cartoons, tables, charts into the text... to break-up the learning into 'short bites'. I really miss useful cartoons and hand-sketch illustrations not present in current-generation documents... they provide insight with humor. These editorial characteristics help readers retain the essence of the information!
That's a good shop briefing. Regarding cold worked holes, ASTM Special Technical Publication 610-EB has a study on how in addition to the beneficial compressive stress at the hole surface of a cold worked hole, there are tensile stresses a short distance away from the hole which may increase stress corrosion susceptibility. They looked at an actual aircraft part failure caused by cold working holes in 7075-T651, then did lab testing with different degrees of interference fit. The bottom line was that you can't blindly go cold working without considering what other factors could come into play. It was suggested that the interference be varied so that the calculated max tensile stress stayed below the SCC threshold stress even if that reduced the theoretical fatigue life improvement.
 

wktaylor

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R.B. ...

Please be very cautious... IF this is the report is "ASTM STP610 Stress Corrosion—New Approaches" it was published in 1976. IF NOT... please provide a link [if possible].

There have been massive technical studies done/published... and real-world applications of FTI-SS-CX... since 1976... by OEMs and individual/organizational researchers. Many early studies, have been refined/over-ridden by later studies and real-world tests and operations. And of course, some hard lessons have been learned in the refinement processes.

NOTE. Alloy '7075 in the -T6511 temper' is exclusive to extruded shapes that have been solution-heat=treated and quenched to attain the unstable 'W' temper... and then stress-relived by stretching/straightening... then age-heated to final -T6xxx temper... which is actually very beneficial OEM 'processing' of the extruded shape. Also, NOTE... in the 1970s there were some VERY serious aluminum alloy heat treating errors made that cost aircraft manufacturers 'hundreds of millions of $$s' to correct. Like I said, a lot of hard lessons learned since the 1970s.

HOWEVER...

R.B. ...May I make a suggestion... let's steer this conversation back to solid rivets... all facets... stay away from fasteners/fastening processes generally reserved for corporate/commercial/MIL aircraft. This discussion path is now 'wayyyyy too technical' for this forum thread. This might be something I'd address in the Eng-Tips forums. NOT here.
 

BJC

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This discussion path is now 'wayyyyy too technical' for this forum thread. This might be something I'd address in the Eng-Tips forums. NOT here.
I’ve enjoyed it so far. Please continue .. waiting to hear about the stress corrosion issues in Lockheed Georgia aircraft.


BJC
 

Rataplan

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I believe these principles also hold equally true with rivet holes. Although most airplanes are designed with a very large safety margin, traditional aircraft rivets are really supposed to be installed in holes that are deburred and reamed. The "dreamer" or drill-reamer is really the best possible tool for this.
Safety Margins are intended for unknown causes not to cover for not applying "in aviation common practises" .
 
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