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6061-T6 1
- Thread starter RoarkS
- Start date
hahahaha - good one - joke of the day i would say. Now i have to explain this to all components i designed for the project spring this year - looks like they are not obeying industry trend - shame of them. About fatigue - this is nothing except truth, ive scrapped 3 bicycle frames during 2 years due to fatigue cracks, all made out of hydroformed and welded 6061 
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.- Thread starter
- #5
It's a riveted/bolted sheetmetal assembly.
He said it should have been 2024-T3.
Holding up probably the smoothest running electric motor pump on the face of the planet between two frames.
Looking at the S-N curves... 6061-t6 looks shifted to the right of 2024-t3 by an entire exponent (if that's how one even says that???)...
In my other life, I was actually admiring some kit planes for using 6061 instead of duralum!
Got me thinking Jaceb, how long do you think a 2024 bike would last? would probably have to be riveted together!
He said it should have been 2024-T3.
Holding up probably the smoothest running electric motor pump on the face of the planet between two frames.
Looking at the S-N curves... 6061-t6 looks shifted to the right of 2024-t3 by an entire exponent (if that's how one even says that???)...
In my other life, I was actually admiring some kit planes for using 6061 instead of duralum!
Got me thinking Jaceb, how long do you think a 2024 bike would last? would probably have to be riveted together!
well, he does have a point; 2024T3 is the typical "go to" for s/m. that said, you can do a fatigue assessment. that said (!?) the DER can make your life a pain ... "prove to me that your spectrum is correct" and other gems. I think the subtext is "I'm not going to (or I don't want to) approve this (without a material change)".
The best argument I think is to show very high static MS, and a very conservative fatigue analysis. If more push back, then a ground test to measure loads in operation (maybe the start up induces higher loads than running loads ?), once you've done that you can easily do a flight test (monitoring the loads in the damn thing in-flight, as if that'll make a difference), the last straw (before changing the material !??) would be an ICA inspection (visual, at C check).
another day in paradise, or is paradise one day closer ?
The best argument I think is to show very high static MS, and a very conservative fatigue analysis. If more push back, then a ground test to measure loads in operation (maybe the start up induces higher loads than running loads ?), once you've done that you can easily do a flight test (monitoring the loads in the damn thing in-flight, as if that'll make a difference), the last straw (before changing the material !??) would be an ICA inspection (visual, at C check).
another day in paradise, or is paradise one day closer ?
I never specify 6061 for any parts, with following exceptions.
Weld-assemblies [6061-T4... then AHT to -T6 after welding*]
Aluminum or magnesium casting replacement... 100% machined.
IF part is to be welded into another Assy, then MF 6061-T4 and AHT to -T6 after welding*
IF part is used 'as-is', then MF 6061-T6
* AHT from -T4 to -T6 is an effective stress relief operation, for the weld-affected zones.
IN a few instances... for both 'light-structural' and non-structural parts... I have specified 6061-T4 or -T6 for flat or straight-bent and mild-compound formed parts.
Regards, Wil Taylor
o Trust - But Verify!
o We believe to be true what we prefer to be true. [Unknown]
o For those who believe, no proof is required; for those who cannot believe, no proof is possible. [variation,Stuart Chase]
o Unfortunately, in science what You 'believe' is irrelevant. ["Orion", Homebuiltairplanes.com forum]
Weld-assemblies [6061-T4... then AHT to -T6 after welding*]
Aluminum or magnesium casting replacement... 100% machined.
IF part is to be welded into another Assy, then MF 6061-T4 and AHT to -T6 after welding*
IF part is used 'as-is', then MF 6061-T6
* AHT from -T4 to -T6 is an effective stress relief operation, for the weld-affected zones.
IN a few instances... for both 'light-structural' and non-structural parts... I have specified 6061-T4 or -T6 for flat or straight-bent and mild-compound formed parts.
Regards, Wil Taylor
o Trust - But Verify!
o We believe to be true what we prefer to be true. [Unknown]
o For those who believe, no proof is required; for those who cannot believe, no proof is possible. [variation,Stuart Chase]
o Unfortunately, in science what You 'believe' is irrelevant. ["Orion", Homebuiltairplanes.com forum]
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1
- #9
LiftDivergence
Aerospace
So you've said the this is a riveted sheet metal assembly "Holding up [an] electric motor pump... between two frames".
OK. So it sounds like this component is only responsible for supporting an item of mass which will see inertial loads from said mass. It doesn't sound like this component supports aerodynamic loads or major pressure loads.
Seems like this might be classified as secondary structure. And damage tolerance analysis is not usually performed for secondary structure. So I have to ask, is this an STC where you are classifying this as primary/secondary and making a determination of whether it is fatigue critical alteration structure? What is driving the need to do a DTE in the first place?
If the loads are really just inertial based on the electric motor mass, the stresses are likely to be very low.
6061-T6 sheet actually has a higher detlaK_th for typical stress ratios. What is the peak stress in your part and can you show the max deltaK is below the threshold?
If you're looking at S-N data you might be able to show the same type of thing for the "endurance limit". But of course, there's always the question of whether the motor induces and vibratory loads, but that depends on the design. You should ask the DER why he feels this should be considered fatigue critical structure. From my perspective, I would be more concerned with the attachment holes you are putting in the frames, and the effect on the continued airworthiness of the OEM frames. Does your part create a hardpoint on the frames? Focusing on the motor mount itself seems odd to me without more detail.
Keep em' Flying
//Fight Corrosion!
OK. So it sounds like this component is only responsible for supporting an item of mass which will see inertial loads from said mass. It doesn't sound like this component supports aerodynamic loads or major pressure loads.
Seems like this might be classified as secondary structure. And damage tolerance analysis is not usually performed for secondary structure. So I have to ask, is this an STC where you are classifying this as primary/secondary and making a determination of whether it is fatigue critical alteration structure? What is driving the need to do a DTE in the first place?
If the loads are really just inertial based on the electric motor mass, the stresses are likely to be very low.
6061-T6 sheet actually has a higher detlaK_th for typical stress ratios. What is the peak stress in your part and can you show the max deltaK is below the threshold?
If you're looking at S-N data you might be able to show the same type of thing for the "endurance limit". But of course, there's always the question of whether the motor induces and vibratory loads, but that depends on the design. You should ask the DER why he feels this should be considered fatigue critical structure. From my perspective, I would be more concerned with the attachment holes you are putting in the frames, and the effect on the continued airworthiness of the OEM frames. Does your part create a hardpoint on the frames? Focusing on the motor mount itself seems odd to me without more detail.
Keep em' Flying
//Fight Corrosion!
I think the DER is thinking it's an odd choice of material for a brkt, which it is.
I think he then said "poor fatigue performance" as giving weight to his (correct) opinion, like I said I think the subtext is "I don't like this" and "I don't want to sign for this".
Once a DER has an opinion in his mind, you have a hard job to change it ! I don't think it's a matter of being rational and showing infinite fatigue life ("I don't believe your spectrum").
And the more you say "but it's just a chicken-shit little brkt", the more he'd say "so it's easy to change".
What's it like to form 6061T6 ? ... as easy as 2024T3 ?
another day in paradise, or is paradise one day closer ?
I think he then said "poor fatigue performance" as giving weight to his (correct) opinion, like I said I think the subtext is "I don't like this" and "I don't want to sign for this".
Once a DER has an opinion in his mind, you have a hard job to change it ! I don't think it's a matter of being rational and showing infinite fatigue life ("I don't believe your spectrum").
And the more you say "but it's just a chicken-shit little brkt", the more he'd say "so it's easy to change".
What's it like to form 6061T6 ? ... as easy as 2024T3 ?
another day in paradise, or is paradise one day closer ?
6061 is easier to form than 2924 if heat treating is involved, due to the fact that 6061 gives you a longer window of time if any subsequent hand straightening operations are involved. Of course the trade off is lower tensile strength with 6061 T6 having a yield point of 40.000 psi, and 2024 T4 having a yield point of 42.000 psi. Of course if you want to take 2024 up to T8 you can get 60.000. If you are just cold forming 6061 T6, it and 2024 T3 . The 2024 requires an inside bend radius of 3T while 6061 T6 can be bent at 1.5 T for 0.063" sheet, so you can see that 2023 requires more care to avoid cracking. Non the less it sounds like the DER is making a molehill out of a trivial matter.
B.E.
You are judged not by what you know, but by what you can do.
B.E.
You are judged not by what you know, but by what you can do.
RoarkS,
Are you designing primary aircraft structure, or appliances mounted inside the aircraft. I have designed all sorts of sheet metal boxes that were mounted inside aircraft, and I specified 5052[‑]H32. This is fabricated reliably by commercial vendors. Structurally, I was concerned about rigidity under vibration. Nobody complained. I was not required to design absolute minimum weight. Aluminium 2024[‑]T4, and 6061[‑]T6 for that matter, are difficult to bend, and should be restricted to qualified fabricators. For commercial fabrication, 5052 is less likely to have cracks in it.
I designed lots of optical devices out of 6061[‑]T6. These had to be rigid, so they saw low stress even under crash conditions.
--
JHG
Are you designing primary aircraft structure, or appliances mounted inside the aircraft. I have designed all sorts of sheet metal boxes that were mounted inside aircraft, and I specified 5052[‑]H32. This is fabricated reliably by commercial vendors. Structurally, I was concerned about rigidity under vibration. Nobody complained. I was not required to design absolute minimum weight. Aluminium 2024[‑]T4, and 6061[‑]T6 for that matter, are difficult to bend, and should be restricted to qualified fabricators. For commercial fabrication, 5052 is less likely to have cracks in it.
I designed lots of optical devices out of 6061[‑]T6. These had to be rigid, so they saw low stress even under crash conditions.
--
JHG
- Thread starter
- #14
Ugh... I wrote a whole paragraph and lost the page.
100% Arizona. Secondary structure. appliance. inertial loads only. If it fails it falls into control cables... If they were not there it would probably punch thru the skin.
I do a lot of 6061-t6 1/8" and cold form bend it at 3T all day long... used to it straight out of AC43.13.
This isn't that thick, but it's certainly thicker than the thickest 2024 I've ever messed with (about .040). If it were welded I would use 5052.. again do that all day long in 1/8" for fuel tanks... but it's not welded so we didn't.
I've stood on the things which equals about 12G static. I'm only required 8G... so we're going down that road too.
100% Arizona. Secondary structure. appliance. inertial loads only. If it fails it falls into control cables... If they were not there it would probably punch thru the skin.
I do a lot of 6061-t6 1/8" and cold form bend it at 3T all day long... used to it straight out of AC43.13.
This isn't that thick, but it's certainly thicker than the thickest 2024 I've ever messed with (about .040). If it were welded I would use 5052.. again do that all day long in 1/8" for fuel tanks... but it's not welded so we didn't.
I've stood on the things which equals about 12G static. I'm only required 8G... so we're going down that road too.
RoarkS - interesting theoretical question - i mean riveted 2024 bicycle frame. If we omit all technological issues related to riveting such tiny frame i would take risk and say that fatigue life of such thing will be less than this one welded out of 6061 - all cracks i had on bicycles were located exactly in the spots close to welds and in heavily loaded parts of frame never on welds itself - everything works according to theory ..
. I think major problem with 6061 frames and welding is technological one - postwelding heat treatment not necessary must be the perfect one. I think steel is still best material for this kind of stuff - i know that it will immediate contra from composite enthusiasts but for me this is material suitable for competition-grade bicycles only - no chance for long life in everyday use environment. Sorry guys for drifting out off aviation issues .
. I think major problem with 6061 frames and welding is technological one - postwelding heat treatment not necessary must be the perfect one. I think steel is still best material for this kind of stuff - i know that it will immediate contra from composite enthusiasts but for me this is material suitable for competition-grade bicycles only - no chance for long life in everyday use environment. Sorry guys for drifting out off aviation issues .
mfgenggear
Aerospace
6061 is a pain to weld, however if the the procedures are correct, and proper NDT it is doable. proper specifications and requirements are required.
if the requirements for the bracket/structure is met then it should be fine. what ever material used. and if the S/N curve is in acceptable limits.
sure there is always a better material but it must have justification.
2 cents
if the requirements for the bracket/structure is met then it should be fine. what ever material used. and if the S/N curve is in acceptable limits.
sure there is always a better material but it must have justification.
2 cents
A whole lotta years ago, I served on a Perry Class frigate. The superstructure was aluminum, I'm pretty sure it was 6061. They were notorious for cracking. I took these pictures after we were caught in a storm for 3 days.
Wouldn't want to see that happen to my airplane.
Brad Waybright
The more you know, the more you know you don't know.
Wouldn't want to see that happen to my airplane.
Brad Waybright
The more you know, the more you know you don't know.
LiftDivergence
Aerospace
drawoh said:
I don't think anyone here is claiming that there is not a safety concern. Secondary structure must be evaluated for damage which may then affect primary structure, change aerodynamic characteristics, or change the overall stiffness. This detail needs to be substantiated just like anything else.
The main thrust of the discussion (at least I thought) was that the approving party seemed to be loath to sign just based on the the fact that the material was not "industry standard". While in fact, proper substantiation could show that it is fine (without much trouble).
The initial post pointed out concerns over fatigue issues. Most metals, including this one, will behave in a brittle fashion eventually under repeated load. But there is some threshold of stress below which this will not happen. Secondary structures generally have very low internal stresses and do not warrant full DT evaluation.
Multiple OEM's definition of Fatigue Critical Baseline structure:
"By definition, does not include miscellaneous secondary structure such as brackets and clips".
Generally FCBS is explicitly defined in structures tables, and something like a motor mount might be classified if it is subject to significant vibration loads or other special considerations, which I noted above.
Even the issue you note above, attachment during a survivable crash, is not a fatigue issue, and would have little to do with cracking, other than from a residual strength or net section yield perspective. That is a static failure issue under the maximum inertial load envelop. So, based on the first post, probably not what the DER was driving at.
This leads me back to my previous post, asking the DER why he feels this should be classified as FCBS, and why a change in material for better damage tolerance properties is required for a bracket. I don't have all the details, so I can't guess. He might have good reasons. But the point is, it should be a discussion with the DER. It is not the job of the DER to impose their will on the design. It is their job to review the design, as is, and either sign that it meets the requirements or not. If you can show it meets the requirements, there shouldn't be an issue, and the DER should be able to tell you the exact regulation he thinks you are not meeting if you use 6061-T6. But I would want more than a simple statement.
He might try to lean on 25.601 (very general). It sounds like this bracket is your own design. Is this an STC? Who is controlling the classification of the structure?
And, after all of this, I still think nobody has discussed the most critical issue, as I brought up above. You are attaching this structure an OEM frame I surmise. If so, you are affecting what is likely a Primary structure, PSE, FCBS. How thick is the bracket compared to the frame? How many fasteners are being used? Is the net area of the frame still acceptable? Does the bracket create a hardpoint on the frame? Have you re-evaluated the rogue flaw capability of the frame?
It is odd to me that for an installation like this, more attention would be paid to the bracket itself than to how the installation affects the existing structure, which is likely to be the critical detail.
Keep em' Flying
//Fight Corrosion!
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