why no Kt at fastener hole for limit load? 1

[AeroNatanio]

Hello,

Sorry if this is a stupid question.

Why, for static calcs of fastener holes @ limit load, do I not use a stress concentration 3.0 around fastener holes?

I understand that at ultimate it all starts redistributing. So no stress concentration would apply there! But why not a stress concentration 3.0 at limit load.

Should it not be the critical load as well? A stress peak factor 3 for holes is more than the factor 1.5 for ultimate. Besides yield stress is lower than ultimate allowable. Should that not make the limit load for all fasteners critical?

Please advise I am thankful

 

[victorzv]

Because there is no sense to do that for a static strength analysis. Plasticity (permanent deformation) is localized in a small area in this case and USUALLY do not affect the function of the part.

Fatigue analysis is a different story.

 

[Stezza]

FAR 25.305 states that "The structure must be able to support limit loads without detrimental permanent deformation"
Generally any plastic region around a fastener hole is small and the stiffness of the surrounding material that has not been yielded forces the plastic material back into its original volume.
Therefore in a global sense the part does not exhibit permanent detrimental deformation.
Thus for parts under axial load the appropriate check is that the average stress at any cross section is below the yield stress, rather than that a localized peak stress is below the yield stress.

 

[AeroNatanio]

Thanks victorzv and Stezza! I really appreciate your experience and the feedback and help.

Okay it does make sense that for static calcs considerations on limit, ultimate, fty and ftu are different and allowed to consider redistribution in plastic range.

Fatigue and static are truly different views. Now I can accept this.

Can I ask another question what happens if it involves fatigue?

Then suppose on a very first cycle, a load (limit * Kt) passing yield stress would that not cause the part then to have a one cycle life. I mean would it initiate a crack right away?

Also if I look up Fty and TYS in ARMMPDS, but it seems to be different values too for the material sheets and for the displayed fatigue chart (example: Fty is 48 ksi and TYS is 54 ksi - unnotched, Ref. 3-71 and 3-119). This may disturb my calc a bit i do not know!

then .. is limit load not a part of a fatigue spectrum in airplanes? I understand the limit load has a very low occurence anyways?

 

[Stezza]

For Static strength analysis you use either the A or B basis allowables - the page 3-71 allowables in your case. For a definition of A and B basis see Section 9.1.6 in the MMPDS. If in doubt use A basis allowables - it can only be conservative.

While I am not 100% sure my guess is the TYS on page 3-119 are reflective of the actual Fty of the fatigue specimens that were tested.

With regard to fatigue, this is a huge and complicated subject and you really need to read some text books and talk to your co-workers.

Firstly crack initiation and part failure are different things, in a fatigue test a part may have a crack but this crack is not yet long enough to cause failure of the part. Failure life is measured by cycles to failure not cycles to crack initiation.

Also an application of limit stress (or over) does not necessarily lead to crack initiation or part failure. You can see from the S/N curves in MMPDS Figure 3.2.3.1.8(g) Kt = 2 that for the upper curve (mean stress 30ksi curve) that if a maximum stress of 60ksi (nett section stress – not local peak stress at the notch) is applied (i.e a stress cycle between 0 and 60 ksi) that you have a fatigue life of 4000 cycles.

 

[rb1957]

a 1 cycle life means the part would have exceeded ultimate strength. you can surely apply an elastic Kt of 3 to a hole and get a low allowable ... your structure will be heavy and strong (or strong and heavy if you prefer).

for fatigue considerations, the local material could well be yielded (i mean it is yielded about the tip of a crack). depending on the spectrum occassional high loads (over-loads) are beneficial to the structure, creating large yield zones that the crack tip has to fight it's way through ('cause when the load is removed at the end of the cycle there is a large compresion stress in the yielded material.

 

[victorzv]

Good questions, AeroNatanio!

And you have already receved very good answers.

I'll just try to put my two cents on the subject.

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Can I ask another question what happens if it involves fatigue?
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Nothing extraordinary, just be careful.

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Then suppose on a very first cycle, a load (limit * Kt) passing yield stress would that not cause the part then to have a one cycle life. I mean would it initiate a crack right away?
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For ductile materials, which aerospace metals are, the answer is NO.

rb1957 has pointed out the beneficial effect of OCCASSIONAL overloads, however the line between "occassional" and "damaging" is not clear outright. You must have a good analysis tool for your spectrum.


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Also if I look up Fty and TYS in ARMMPDS, but it seems to be different values too for the material sheets and for the displayed fatigue chart (example: Fty is 48 ksi and TYS is 54 ksi - unnotched, Ref. 3-71 and 3-119). This may disturb my calc a bit i do not know!
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Stezza provided excellent explanation.
Fty is a safe value for static analysis. TYS on the S-N curves are an actual average (50% probability) value for the tested specimens. This value should be used in a fatigue analysis.


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then .. is limit load not a part of a fatigue spectrum in airplanes? I understand the limit load has a very low occurence anyways?
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I saw a thread on this subject. There were some good ideas.

My point of view is as follows.

In terms of paragraph 25.305, i.e. static strength, it can.
(However, there is a minor contradiction between "static strength" and "fatigue spectrum").
But it is proven bad practice to encounter loads close to the limit one often. If you know that you will reach high g in every second flight (aerobatic) then set your limit loads at, say, 20 g.


There is an equivalent of the limit load in fatigue - more precisely Damage Tolerance - analysis (FAR 25.571). It is used to determine residual strength of the part. This pretty much a special question, so that it is better to refer directly to MIL-A-83444 or JSSG-2006.