Fuselage Stringers 2

[davidnoc]

Hi All,
I do not have much experience with FEA.But I am trying to output stringer (axial) loads from Boeings in-house FEADMS tool. I observed actual drawing area for Stringer in the crown has lesser area than the FEADMS model area for one of the stringer.

With this load and actual drawing area the stringer can not make positive margin.

So can I reduce the load obtained for this stringer considering the actual area based on area propornality? Is it linear? Or the load obtained for the stringer is stil correct, assuming that it does not depend on area.

So I just want to know whether my output for stringer loads are right or wrong, as the area for one of the stringer entered in model is more than actual.

Thanks

 

[rb1957]

depends on how it's failing ... long column or short (crippling). unless the true area is Way more than the modelled area the load in the stringer won't change much.

is there a stringer pad you can take advantage of ? don't like the idea of redistributing load into the skin if the stringer's buckling.

btw watch you spelling ... it's proportionality not "propornality"

 

[YoungTurk]

It is likely the element area is increased to account for some section of skin which it is attached to. This is referred to as the "effective width" of skin and is generally taken as 20-30 times the skin thickness.

 

[rb1957]

it could be, but that'd be pretty crude for the largest (ok, one of the largest) commercial airplane designers and manufacturers. it'd also mean that the skin is modelled with shear effective elements only. it'd also mean a great deal of post-processing to understand the loading of the skins.

it could be that effective skin was mistakenly added into the stringer element ... ie the effective area was added to the stringer, and membrane effective skins were also used for the fuselage (effectively double counting the skin).

 

[davidnoc]

Hi all,

Actually the stringer pad up thickness is considered in the FEA model thats why the model has higher area(80 % more than the drawing, without padup).

The stringer is failing in net tension check. In my calculation I am considering the area of stringer without padup.

Since the padup is local I thought may be I can reduce the load obtained from model based on drawing area as it may be way higher due to larger area(I think as the stiffness is increased the load on this stringer is increased)

Let me know how you guys think.

Thanks

 

[rb1957]

ok, there's a padup on the skin, incorporated inot the stringer element.

1st, the stringer ... reduce the load in the stringer by the area ratio (actual stringer area/ FEM atringer area), having verified that the FEM stringer area = actual stringer area + skin pad-up area.

then the pad-up ... consider the pad working with the skin ... the pad could be working at a higher stress than the skin panel, or you could rationalise that both should be working at the same stress level. then consider the loadpath into (and out of) the pad up ... the shear conntection is obviously carrying an increased load at the pad up.

finally, you've got a tension problem ?? that tells me that Boeing has given you reduced tension allowables to avoid fatigue problems. I consider this to be a "soft" failure ... one that would produce required inspection rather than one tha requires redesign (ie like a compression failure would).

good luck

 

[davidnoc]

So is this the high stiffness which has resulted in higher load?

 

[rb1957]

unlikely ... how significant is the pad up compared to the gross area of the fuselage ? maybe there is a large load input near this site ?

i suspect the pad up is there so that they can use a standard stringer extrusion, found an overload, and wacked on some more A to deal with it.

 

[SWComposites]

Some questions:
- is this a metal or composite fuselage?
- is this for a new design or a modification/repair?
- have you asked the question to the Boeing stress lead for the fuselage section?

 

[40818]

I would take a quick step back with this one and have a more global view.

Your stringer is situated on the (upper) crown of the fuselage and is under a tension load, hence it is fuse bending case. The strain distribution of the fuselage is putting the top into tension and bottom into compression. The strain distribution is governed by the whole fuse section, not a small bit of area difference of the stringer. You cannot simply take an area ratio of the FEM & actual and manipulate the stringer stresses/loads. Put simply, the local change in stringer area will probably have only a negligible change in the stringer true stress.

Your nett tension problem....
1) Are you increasing the stringer stress by the nett-gross area?
2) Are you reducing the material Ftu by accounting for holes etc (i.e Hawker Siddley Aircraft handbook)
3)Have you thought about redistribution effects