777 Floor Panel Stress Analysis 4

[audacious1234]

Hi,
I am very green to this stress analysis field and I am hoping to get guidance from this forum.
I am doing 777 floor panel strength substantiation. Floor panel material BMS4-20 Type III. Loading is console load from seats which in Z direction. For conservative analysis I am considering it as a panel with all edges simply supported and load applied in middle. But I couldnt find any formula or theory to calculate bending moment/ bending stress. Does anyone of you know how to do this? or Do you know any other methodology to do strength substantiation of these kind of panles?
Thanks

 

[MiketheEngineer]

Talk to your boss or other seasoned engineer. Since I will surely fly on this plane - lets get it right.

There is an old story that while the 747 floor was fine - women with high heels would destroy them. Keep that in mind.

 

[audacious1234]

That is what I am doing and asking seasoned engineer on this forum. Dont worry my approach is very conservative.

 

[GBor]

Find a similar structural substantiation report and model your approach after it. Another option is to contact a DER and ask them for any information they may be able to provide. If you are doing this for a particular application, see if your company can get the current structural substantiation for 777 from Boeing. There are several other avenues, but these are the first few that come to mind.



Garland E. Borowski, PE
Star Aviation

 

[FieldTeam]

off subject:

"There is an old story that while the 747 floor was fine - women with high heels would destroy them...."

Boeing made their improved floor panels, BMS4-23, to better handle in service wear (impact and rolling load).
BMS4-20 is graphite epoxy face with nomex core.
BMS4-23 is fiberglass flame retardent epoxy face with aluminum honeycomb core.

One of my co-workers said that the 777 has composite graphite floor beams with BMS4-23 floor panels. Maybe the type floor panels are a customer option.

 

[audacious1234]

Thanks for all the replies but I didn’t get an answer related to my technical question. How should I calculate bending moment for a honeycomb panel all edges of the panels are simply supported? (E is not known so I can’t use Roark’s formulas for plates with all edges simply supported?).
Thanks

 

[hokie66]

You won't be able to gain an education in structural analysis in this forum. E does not affect bending moments.

 

[audacious1234]

Thanks for your reply.
I know E doesnt belong to bending moment. But it is belong to bending stress and in Roark you will get bending stress not bending moment.

Regards!

 

[hokie66]

Well, I am sorry, you are wrong again. If you know the bending moment and the moment of inertia, you know the bending stress, regardless of the material properties.

 

[audacious1234]

Seems like you dont have Roark. I didnt say I know bending moment I said I am looking for a way to calculate bending moment and I cant use Roark formula because it needs E value to get bending stress and thats the only formula I have. Roark doesnt have anything for bending moment.
Thanks a lot again. Again the problem is how one should calculate bending moment of the panel which is simply supported on all four edges and loading is concentrated load in middle.

 

[hokie66]

It may help to post the formula for stress you are using. One of us is hopelessly wrong. Are you sure the formula with E is not to determine deflection?

 

[inertia4u]

A plate s-s on all sides is a statically indeterminate structure - which means that in order to solve bending stresses and moments, you need to to account for stiffnesses (or bending rigidity - i.e. deflections) of the plate in various directions. So I can see young's modulus entering into the equation.

I would think that you could conservatively reduce this to a 2D beam problem. Assume a unit width (in the short direction), and then treat the plate as a s-s beam, length = long dimension, Upper face sheet = Upper chord, Lower face sheet = lower chord, and core is the shear web. Calculate the max moment for this beam (where the load is applied). \

Then assume that the face-sheets carry axial load only, and the core carries shear only.

-----
Nert

 

[GBor]

Do you have access to the Boeing Design Manuals? The 6700 series manuals cover sandwich panels with BDM 6710 looking like a good candidate for you. without some significant information, it would be difficult to actually calculate the numbers for you.

 

[RPstress]

Roark Ed 6 table 26 case 1b gives max stress for uniform plates with a central load. You can convert Roark's formula for stress to one for moment. For a plate this is done by multiplying the stress formula by t^2/6. The resulting formula gives the moment per unit width at the plate centre. Basic sandwich panel formulas can then be used to find the max stress. Note that the shear distribution near the load cannot be addressed with Roark. Also, to use Roark for the deflection of a sandwich panel it's necessary to find the equivalent plate thickness for the sandwich panel.

NB: for uniform isotropic plates you don't need E to find moments/stresses. The *relative* stiffness along and across the plate (and indeed the ratios of those to the plate torsional stiffness, provided nu is about 0.3) is unchanged by E.

Note that if the sandwich panel face sheets are not quasi-isotropic then the Roark plate analysis is wrong. In this instance the Es *are* necessary to find the moment distribution/stress. Also, if you get into the detail then the core shear stiffness will also change the bending moments (even with isotropic face sheets), because the shear modulus along the ribbon direction is more than it is across the ribbon (the core being honeycomb). The plate stiffness is the inverse of the sum of the bending and shear flexibilties.

Audacious: you should be being supervised by someone knowledgeable who can give this sort of advice while you build up your own basic knowledge. As GBor says, a good set of static stressing manuals would help. Roark is good, but it is not the be all and end all.