As a first post in a very interesting forum, I would like to ask 3rd party repair engineers what they use as an ultimate static stress (tensile) when designing “scab patches” on the fuselage skin. I think using Ftu leads to an over dimensioned repair when you consider that the fuselage is basically a thin walled tube in bending which fails in compression and leaves lots of opportunity to allow for local yielding in the upper shell. I think Fty is more appropriate. (reasons will come later)
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Ultimate stress in a pressurized fuselage 2
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astroclone
Aerospace
As a repair engineer we make our repairs as conservative as possible.
Usually 1 gage heavier and based on Ftu. We don't have access to the stress reports so we assume that the worst case scenario would be Ftu failure.
The fuselage by itself would fail in compression, that's what the stringers and frames are for.
T
Usually 1 gage heavier and based on Ftu. We don't have access to the stress reports so we assume that the worst case scenario would be Ftu failure.
The fuselage by itself would fail in compression, that's what the stringers and frames are for.
T
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- #3
Thanks astroclone
That is just my point, if I design to Ftu I will more than often require more than the 3 fastener rows that are shown in Airbus and Boeing SRM’s, going 1 gauge thicker means I have a larger bending offset in the fastener and this leads to a reduction in the fatigue life. As you say the fuselage fails in compression, these compression stresses do not approach the order of magnitude of Ftu.
As a matter of interest, do you have more experience with Boeing or Aus aircraft?
That is just my point, if I design to Ftu I will more than often require more than the 3 fastener rows that are shown in Airbus and Boeing SRM’s, going 1 gauge thicker means I have a larger bending offset in the fastener and this leads to a reduction in the fatigue life. As you say the fuselage fails in compression, these compression stresses do not approach the order of magnitude of Ftu.
As a matter of interest, do you have more experience with Boeing or Aus aircraft?
astroclone
Aerospace
No experience with Airbus.
Lots with Boeing, MD, Sabreliner, Hawker, Bombardier, etc.
If the doubler is too thick you can use more than one thickness and stagger the last fastener rows.
Or taper the edges so the last fastener row is going through thinner sheet. Etc...
T
Lots with Boeing, MD, Sabreliner, Hawker, Bombardier, etc.
If the doubler is too thick you can use more than one thickness and stagger the last fastener rows.
Or taper the edges so the last fastener row is going through thinner sheet. Etc...
T
I have around 7 years of experience designing and analyzing Boeing 737 aircraft- having been a fleet support stress engineer for the big B. I can tell you that the without any load data, we sized the repair to ultimate capability (Ftu), and we will review the repair from a fatigue standpoint as well.
When load data is available, we will size to that, however, from what I remember, fatigue would sometimes drive the need for an additional fastener row (i.e. taking it out to 4 rows, such as astroclone said).
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Nert
When load data is available, we will size to that, however, from what I remember, fatigue would sometimes drive the need for an additional fastener row (i.e. taking it out to 4 rows, such as astroclone said).
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Nert
- Thread starter
- #6
Thanks for both your replies astroclone and inertia4u but I don’t want to have to design always on the conservative side, sometimes it is not possible to put in that 4th fastener (the fatigue guys I’ve met tell me, that 4th fastener wont help him much anyway. I have some Airbus RAG curves that show this if any of you are interested) row due to spacing and when I start telling the mechanic he’s to machine a “finger” doubler or put a step at the last fastener row he’ll flip out. I can give 3 reasons why Ftu is way too high:
1. The reason in the first post, local yielding in the thin walled tube (fuselage) will not permit the tension stress to go beyond Fty, it will always find a way to include more effective tension area.
2. Modern day jet liners (including those using Glare) are designed to a (undisturbed area) fatigue life of 90 to 110 N/mm² fatigue stress converting this to an ultimate static stress (Nz = 2,5 and 1,5 delta P) you obtain a stress round about yield.
3. (maybe a little weak but) Linear FEM models do not obviously include the inelastic behaviour of the structure, reducing the ultimate elastic stress (Ftu value) using Ramberg-Osgood you get a value under Fty when considering 2024 T3 properties.
Is there a DER here that won’t sign off a repair when I give him these arguments?
I can give you a reason for not using Ftu: when I put holes into the structure for the fasteners, I have reduced the net section tension area of the original and so I get automatically a MoS less than 0 and a reserve less than 1.
Mike
PS sorry if I’ve went on a bit.
1. The reason in the first post, local yielding in the thin walled tube (fuselage) will not permit the tension stress to go beyond Fty, it will always find a way to include more effective tension area.
2. Modern day jet liners (including those using Glare) are designed to a (undisturbed area) fatigue life of 90 to 110 N/mm² fatigue stress converting this to an ultimate static stress (Nz = 2,5 and 1,5 delta P) you obtain a stress round about yield.
3. (maybe a little weak but) Linear FEM models do not obviously include the inelastic behaviour of the structure, reducing the ultimate elastic stress (Ftu value) using Ramberg-Osgood you get a value under Fty when considering 2024 T3 properties.
Is there a DER here that won’t sign off a repair when I give him these arguments?
I can give you a reason for not using Ftu: when I put holes into the structure for the fasteners, I have reduced the net section tension area of the original and so I get automatically a MoS less than 0 and a reserve less than 1.
Mike
PS sorry if I’ve went on a bit.
When all the joints in the skin of the aircraft have only 2 rows of countersunk fasteners, as there are on most aircraft I work on (light and commuter-size twins), justifying 3 or 4 rows of fasteners is very difficult.
When designing a repair, I question the assumptions in the analysis more often than reaching for a third row. I usually try to show that the skin stress in the repair / antenna installation has been reduced below 12 ksi or so (85 N/mm[sup]2[/sup]).
85 N/mm[sup]2[/sup] = 12 ksi, and 12 ksi * 0.032" sheet = 384 lb/in. Two rows of MS20470AD4's are sufficient.
Sometimes the stress concentrations around repair cut-outs are staggering, and then the design gets "rivety".
I haven't seen any airframe techs "flip out" yet, and the repairs do get signed out.
BTW, why would anybody want to machine a thin edge on a sheet? Bond two thin ones together, and one is longer and wider than the other, to get the same effect, right?
STF
When designing a repair, I question the assumptions in the analysis more often than reaching for a third row. I usually try to show that the skin stress in the repair / antenna installation has been reduced below 12 ksi or so (85 N/mm[sup]2[/sup]).
85 N/mm[sup]2[/sup] = 12 ksi, and 12 ksi * 0.032" sheet = 384 lb/in. Two rows of MS20470AD4's are sufficient.
Sometimes the stress concentrations around repair cut-outs are staggering, and then the design gets "rivety".
I haven't seen any airframe techs "flip out" yet, and the repairs do get signed out.
BTW, why would anybody want to machine a thin edge on a sheet? Bond two thin ones together, and one is longer and wider than the other, to get the same effect, right?
STF
When doing repairs, the best guide is the SRM. If that fails, a good guide is the local rivet patterns. Determine the ultimate load for the longitudinal and hoop dirction joints. Utlimately though, it will be the fatigue and DT issues that drive the design. Static analysis alone is not a sufficiant basis for substantiating a repair for these aircraft.
Nigel Waterhouse & Associates
Aeronautical Consulting Engineers
Transport Canada and F.A.A approval & certification of fixed and rotor wing aircraft alterations: Structures, Systems, Powerplants and electrical. FAA PMA, TC PDA.
n_a_waterhouse@hotmail.com
Nigel Waterhouse & Associates
Aeronautical Consulting Engineers
Transport Canada and F.A.A approval & certification of fixed and rotor wing aircraft alterations: Structures, Systems, Powerplants and electrical. FAA PMA, TC PDA.
n_a_waterhouse@hotmail.com
It was mentioned that the fuselage will fail in compression.
Are we saying that the fuselage will fail in compression at the lower lobe due to bending?
Why won't the fuselage fail in tension on the upper lobe/crown area? Likewise, why won't the fuselage fail in tension in the hoop (circumferential) direction due to
pressurization "fatigue"?
Thanks,
Alex
Are we saying that the fuselage will fail in compression at the lower lobe due to bending?
Why won't the fuselage fail in tension on the upper lobe/crown area? Likewise, why won't the fuselage fail in tension in the hoop (circumferential) direction due to
pressurization "fatigue"?
Thanks,
Alex
astroclone makes it sound like a fuselage will fail in compression. Presumably this could happen at the belly when ultimate tail load and vertical inertia are combined in, say, a high-G load pull-up maneuver recovering from a dive. The skin panels wrinkle and the stringers buckle. The likelihood of this happening sounds remote to me.
I've seen a few failures of fuselage skins in my brief time as an engineer (4 years). All of them have been due to loads normal to the skin that had nothing to do with the original type design. Either it was from a mechanic sitting on the nose as he replaced a windshield panel, or from floppy antennas that had no business being installed on an aircraft,![[mad]](/data/assets/smilies/mad.gif "[mad] [mad]")
or from an inspection plate that wasn't secured properly (tore off in flight, taking skin with it).
STF
I've seen a few failures of fuselage skins in my brief time as an engineer (4 years). All of them have been due to loads normal to the skin that had nothing to do with the original type design. Either it was from a mechanic sitting on the nose as he replaced a windshield panel, or from floppy antennas that had no business being installed on an aircraft,
or from an inspection plate that wasn't secured properly (tore off in flight, taking skin with it).STF
I've seen a fuselage fail in compression! Although it was no design condition. The mechanics raised the nose gear while the airplane was parked and it buckled the crown skin and stringers.
I agree with Nigel. The least conservative thing you can do and get away with is copy a local fastener pattern. If you're repairing an airplane and you don't have access to the stress notes, don't make the repair weaker than the original design.
I agree with Nigel. The least conservative thing you can do and get away with is copy a local fastener pattern. If you're repairing an airplane and you don't have access to the stress notes, don't make the repair weaker than the original design.
Philcondit:
I was reading our 717 SRM on how to size say a rectangular repair doubler for skin that was trimmed out or blended out.
I am wondering if you find the method too conservative:
Take the case where you have excessive blendout due to impact damage. The skin that's blended out has dimensions L x W x T where T=max. thickness of skin blendout.
To size the doubler itself, you pick a material and gage whose Ftu*t of the doubler is greater than Ftu*t of original skin. This ensures that the doubler can carry the original factory skin tensile loads.
Then, you find out how many fasteners you need on say, the "L" edges (two edges).
# of fasteners along "L" edge = (Ftu * L * T *1.15)/(joint allowable found in MIL-HDBK5)
Repeat the same thing along the "W" edges by replacing "L" with "W" above.
Now, you've sized your fasteners to be able to transfer ultimate loads (in addition to a fitting factor of 15%).
Do you think this method is too conservative? Or must an airline repair engineer abide by it at all times when sizing a a doubler?
Frankly, this method seems like the single one any repair engineer would ever need (with no fatigue considerations) at their job, even though it may seem rather conservative.
I was reading our 717 SRM on how to size say a rectangular repair doubler for skin that was trimmed out or blended out.
I am wondering if you find the method too conservative:
Take the case where you have excessive blendout due to impact damage. The skin that's blended out has dimensions L x W x T where T=max. thickness of skin blendout.
To size the doubler itself, you pick a material and gage whose Ftu*t of the doubler is greater than Ftu*t of original skin. This ensures that the doubler can carry the original factory skin tensile loads.
Then, you find out how many fasteners you need on say, the "L" edges (two edges).
# of fasteners along "L" edge = (Ftu * L * T *1.15)/(joint allowable found in MIL-HDBK5)
Repeat the same thing along the "W" edges by replacing "L" with "W" above.
Now, you've sized your fasteners to be able to transfer ultimate loads (in addition to a fitting factor of 15%).
Do you think this method is too conservative? Or must an airline repair engineer abide by it at all times when sizing a a doubler?
Frankly, this method seems like the single one any repair engineer would ever need (with no fatigue considerations) at their job, even though it may seem rather conservative.
astroclone
Aerospace
I was thinking what Sparweb was thinking, too. You don't want too thick of a sheet. So if you have an 0.080 thick doubler requirement use one sheet of 0.032 and one of 0.050.
It's a lot easier to make 2 doublers than to machine steps and tapers and fingers and so on.
As a repair engineer I can't justify doing anything less than an Ftu type repair as stated above. We don't know the loads and manufacturers have a death grip on their stress reports.
Our company even used to make corporate jets and when we repair our own airplanes they fight us for the stress report.
I did manage to do a cool crippling analysis one time, though using the company stress report.
Good times,
T
It's a lot easier to make 2 doublers than to machine steps and tapers and fingers and so on.
As a repair engineer I can't justify doing anything less than an Ftu type repair as stated above. We don't know the loads and manufacturers have a death grip on their stress reports.
Our company even used to make corporate jets and when we repair our own airplanes they fight us for the stress report.
I did manage to do a cool crippling analysis one time, though using the company stress report.
Good times,
T
Koopas:
Your method is certainly the most conservative way to repair: w = Ftu*t (where w is running load in lbs/in). The 15% "fitting factor" may be a little over the top. We normally only apply a fitting factor to single pin joints.
The second method which Nigel alluded to and I commented earlier, requires you to read the airplane's structure. By analyzing the strength of a local, original joint, w = (Pall x # of rows)/spacing, you may find some relief in sizing your repair. As stated earlier, you need to read the structure and pick the right joint.
There are a lot of good posts on this thread about many other things that should be considered when designing a repair. Since you're dealing with a 717, Boeing Customer Service has a checklist for engineers doing repairs on their web site.
Your method is certainly the most conservative way to repair: w = Ftu*t (where w is running load in lbs/in). The 15% "fitting factor" may be a little over the top. We normally only apply a fitting factor to single pin joints.
The second method which Nigel alluded to and I commented earlier, requires you to read the airplane's structure. By analyzing the strength of a local, original joint, w = (Pall x # of rows)/spacing, you may find some relief in sizing your repair. As stated earlier, you need to read the structure and pick the right joint.
There are a lot of good posts on this thread about many other things that should be considered when designing a repair. Since you're dealing with a 717, Boeing Customer Service has a checklist for engineers doing repairs on their web site.
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- #16
Sorry, but I am still not convinced that using Ftu is the only way and most of the posts do say that they find the method conservative. Regarding using 2 doublers instead of 1 with steps etc, you have to remember the 2 doubler solution has a different bending stiffness than 1 thick doubler (different fatigue behavior) and to avoid knife edge effects I am stuck with using a minimum sheet thickness of 0.063” when installing a 3/16” fastener which tends to make the repair doubler overall thicker than I would require, which brings in the offset (fatigue) problem again.
“Reading” the airplane structure and copying what’s there is fine, but again you have to be convinced that the manufacturer has put the joint at the most highly stressed part of the aircraft. Copying and adapting the SRM is the better solution but unfortunately the repair you require just happens to be missing most of the time.
I think I’ll continue using Fty to size the repair.
“Reading” the airplane structure and copying what’s there is fine, but again you have to be convinced that the manufacturer has put the joint at the most highly stressed part of the aircraft. Copying and adapting the SRM is the better solution but unfortunately the repair you require just happens to be missing most of the time.
I think I’ll continue using Fty to size the repair.
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Forgive me if I'm wrong, but there seems to be a little bit of confusion between ultimate and limit loads.
At Boeing, we sized structure to ultimate loads (1.5*limit) or whatever the Boeing requirement was for static analysis. For fatigue analysis, we brought the load down to limit if we did not run fatigue loads.
If you do not have access to loads, then you must size to the capability of the structure, which I agree, tends to be conservative. However, an aircraft rarely "operates" at ultimate load (if ever) so one could make the argument to use Fty since this is closer to the "operational" loads.
To conclude, at Boeing, we used the most conservative analysis possible unless weight became an issue. That way, you knew the structure was good.
Hombre
At Boeing, we sized structure to ultimate loads (1.5*limit) or whatever the Boeing requirement was for static analysis. For fatigue analysis, we brought the load down to limit if we did not run fatigue loads.
If you do not have access to loads, then you must size to the capability of the structure, which I agree, tends to be conservative. However, an aircraft rarely "operates" at ultimate load (if ever) so one could make the argument to use Fty since this is closer to the "operational" loads.
To conclude, at Boeing, we used the most conservative analysis possible unless weight became an issue. That way, you knew the structure was good.
Hombre
- Thread starter
- #19
I agree 1.5 times the limit load is the ultimate load in the aircraft. What I am saying is that the stress produced from 1.5 times the limit load will not exceed the material Fty if we are talking about the fuselage (thin walled tube) structure. See reasons given above.
I am also talking about doing the static analysis. Without knowing the actual loading it is impossible to determine the fatigue life or inspection interval. What you can do however, is, by following good design practise, ensure that the repair doubler is optimised from a fatigue point of view, e.g. do not end on existing fastener positions, avoid installing too thick a doubler, take the doubler all the way around the corner and so on
I am also talking about doing the static analysis. Without knowing the actual loading it is impossible to determine the fatigue life or inspection interval. What you can do however, is, by following good design practise, ensure that the repair doubler is optimised from a fatigue point of view, e.g. do not end on existing fastener positions, avoid installing too thick a doubler, take the doubler all the way around the corner and so on
HI ALL
What interesting forum!
Fristly you always should desing structural repairs acc.
SRM or following mfg. indications. it,s the best method,and the safer too.
The first question of sheet metal guys is why they put only two rows and we,ve to instal 3 or more. Well due cheminal milling , skin panel has different thicknes you
know.the joint has the thicker one you can install 3/16
rivets in.080 TH. sheet but never below .040 all my
handboock tell the same.
Perhaps you can desing a doubler made of two different
thicknes sheet based in Ftu but take acount the knife effect between skin TH. and diameter of rivet installed
escuse my basic English. I,m trying to improve it.
)
What interesting forum!
Fristly you always should desing structural repairs acc.
SRM or following mfg. indications. it,s the best method,and the safer too.
The first question of sheet metal guys is why they put only two rows and we,ve to instal 3 or more. Well due cheminal milling , skin panel has different thicknes you
know.the joint has the thicker one you can install 3/16
rivets in.080 TH. sheet but never below .040 all my
handboock tell the same.
Perhaps you can desing a doubler made of two different
thicknes sheet based in Ftu but take acount the knife effect between skin TH. and diameter of rivet installed
escuse my basic English. I,m trying to improve it.
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