Composite Anisotropy

[Casilero]

A question for those interested in composite panel buckling.

It has been a curiosity as to why edge shear on a typical panel or structure is generally treated as a positive value or magnitude. It seems to be a practice and for metallics, it would be reasonable to assume that as metal is more a less isotropic, then behaviour is reversible. However, a panel subject to diagonal tension across one diagonal caused by shear alone, then reverse the shear direction (e.g. a wing in twist reversal), the buckling mode can differ and the buckling and post buckling capability can be quite different. Sol 106 Buckling I think looks after all this I think in reality. But in the pencil days , it did not.

Add to this, composites as a material. The sensitivity to the direction of shear could be quite profound.

Recently while bench mark testing a new buckling and laminate strength analysis program, I was given the task of verifying the tools simulation results against physical tests conducted by NASA in the 60's as sort of bench mark studies, for balanced composites and unbalanced composites and single curvature panels. The tool did not accept the negative shear value to represent load reversal but the geometry of the test simulation could be reversed so the loads were for all effective purposes be reversed. The result were some what unexpected and might have implications for composites. It was possible to see major differences in behaviour in panel buckling capability especially for low ply numbers. Even some high ply layups up around 20 ply, ocassionally there were substantial differences.

There is a tendency when setting up the loads going into sol 106 to simply put in the magnitude of the shear and not the direction. Sol 106 will take negative shear.

Has anybody actually encountered practical implications from this simplification of shear direction and buckling.

 

[rb1957]

IDK enough to be very sensible but it sounds like something to me. Can you play with the material directions to effectively flip the shear direction ?

 

[SWComposites]

composites data from the 1960's is likely suspect, at best.

NASA published lots of composite buckling reports in the 1990's and 2000's.

Yes, you are correct, the sign of the shear load matters for a composite laminate using tape materials, because the bending stiffness is not the same in the +45 and -45 directions. If you use FEM, you need to apply both shear load directions. Or possibly extract both +ve and -ve eigenvalues.

The composite panel buckling tool that I wrote decades ago calculated separate buckling load for + and - shear.

 

[sdm919]

Yes, and what you have noticed is what makes the difference between "orthotropic" and "anisotropic" analysis. In the bending stiffness matrix, the bend-twist coupling terms D16 and D26 make the difference. When D16 and D26 are not zero, the D matrix is "fully populated" and referred to as anisotropic. When D16=D26=0, the behavior is called orthotropic (or "specially orthotropic"). Many composite analysis tools and handbook formulas are based on the orthotropic assumption because solving the fully anisotropic problem is more difficult. I presume the manual for the tool you are using should mention that when it describes its theoretical basis. Even if it is only a brief description, the term "orthotropic" plate theory would be a dead giveaway that it is neglecting the D16 and D26 terms in the buckling analysis.

Your studies have also revealed when these effects might be important. In a symmetric 4-ply laminate consisting of (+45/-45/-45/+45) they will be significant. But in a 24-ply symmetric laminate with [ (+45/-45)12 ]sym they will be much smaller. On the other hand, the effects will be significant for a 24-ply symmetric laminate with all the +45 and all the -45 plies stacked together (that is: [ (+45)12 / (-45)12 ]sym. That is another reason that it is good practice to disperse the different ply angles rather than clump like plies together.

To get an feeling of how different layups affect the magnitude of these terms, all you need to do is to compute the ABD matrices for a variety of laminates, you don't need to do the full FEA run. Look at the D16 and D26 terms and see how large they are in relation to D11 and D22. Michael P. Nemeth (NASA) uses two non-dimensional "anisotropic parameter" called gamma and delta to quantify these effects. (e.g. see his paper "Importance of Anisotropy on Buckling of Compression-Loaded Symmetric Composite Plates" AIAA Journal, Vol 24, No 11, Nov 1986. although has has many similar papers on the NASA Technical Report Server).

Robert M. Jones also has some discussion along these lines in his book "Mechanics of Composite Materials". See Chapter 5, Bending, Buckling, and Vibration of Laminated Plates.

 

[sdm919]

one more thing I should have added...

An for any given laminate, if you interchange the +45 and -45 plies, you will change the sign of the D16 and D26 terms, but not the magnitude. Again, looking at the ADB matrices of lots of sample laminates should be instructive.