Tim,
Thanks! I recognize you put some thought and effort into this. That is very much appreciated.
You’re a few steps ahead of me. Please help me catch up? Can we start with the panel layup and how you derived the user defined material specifications for FinSim? Did you use “The Laminator” or some other method? If you did use The Laminator are you sure it applies to sandwich panels?
Feckless
Feckless,
Let me supply some of this answer and I'm sure there will be more questions which we can address. I want to preface this by saying I am not a materials or structural or mechanical engineer. My use and interpretation of modeling data and software has been acquired over hours/days, dare I say months of study Federal/DoD and various university/industry docs and texts purchased. Saying that, I am far from formally trained and I may absolutely be wrong in my projections and conclusions--caveat emptor. Now, if you are scared enough, I will proceed (with caution)...
Yes, I did use
The Laminator and the product specs you provided for the core material. The product specs provided just covered the strength specs and not the physical/mechanical specs. So I supplemented the strength specs with the physical properties of light Balsa found in a number of government/forest industry reference documents. The aggregate specs that were used are in the file below:
View attachment EndGrainBalsa.docx
From that point I built a user-defined laminate material aptly named
End Grain Balsa. It was the core of a stack up that was surrounded by equivalent of T300-type carbon fiber / epoxy matrix. I progressively added CF layers and simulated (see pic below, core material not shown). I then subsequently fed the aggregate laminate properties into
FinSim with the appropriate fin geometry given in your Rocksim file and fin thickness as defined by the layup. Summary results were given previously, but I had to get up to 2 layers of CF on each side (5 layers total) to get above Mach 1.1/1.2 on flutter and divergence velocities. Seven layers total provided 15% margin, nine layers total 25% margin. In addition, I did try a few fiber cross-ply variations on the seven and nine layer stacks to provide optimum results. As expected it is the cross-ply stack shown in the previous post. Of course, in addition, in each layering step the flutter results benefited both by a stiffer skin and an overall thicker fin.
For completeness resultant aggregate properties of the 7 and 9 layer stack, respectively, are given below (i.e.,
The Laminator output).
View attachment Laminate15incher_v2.txt View attachment Laminate15incher_v5.txt
Now for your last question... I believe the real question is, "Can The Laminator software adequately model a composite structure that is "dominated" by a vertical honeycomb structure?" My answer is, "I'm not sure." To defend that answer, I will post an excerpt from a well written paper on honeycomb material characterization (Section 4.1.2, Macromodeling, apparently a Master's Thesis from a student at the Imzar Institute of Technology in Turkey). It basically summarizes what is true of laminate analysis in general, it is a simplification and averaging of the characteristics of each of the composite materials, not only in their individual layers, but also across the whole 3D structure. Under that definition, yes, I feel the results are reasonable although I do suspect that the Angle-of-Attack characteristics are conservative (but that maybe should be left for another post). Saying the previous, it is clear that a honeycomb structure with its fiber strength in the vertical direction is a different beast than carbon or fiberglass cloth. At
this link is the full paper to show you how they modeled the subcomponents of each of the honeycombs and the facesheets. So are we close, I believe so. Are we exact, HECK NO!
The macromechanical approach is concerned with the contributions of each ply
to the overall properties, therefore the properties of the fiber and matrix are averaged to
produce a set of homogenous, orthotropic properties. In the case of composite laminate
there is an additional level of complication which arises as a result of stacking several
layers of composites with different orientation and properties. For a given stacking
sequence, the stress-strain relations of a composite laminate can be derived and the
various coupling mechanisms between in-plane and out of plane deformation modes can
be explored. In macromechanical modeling, prediction of failure of a unidirectional
fiber reinforced composite is usually accomplished by comparing some functions of the
overall stresses or strains to material strength limits. Several failure criteria such as
maximum stress, maximum strain, Tsai-Hill, Tsai-Wu have been suggested to predict
the failure. These criteria are based on the average composite stress strain states.
Macromechanical modeling does not consider the distinctive behavior of the fiber and
matrix materials. Although the macromechanical approach has the advantage of
simplicity, it is not possible to identify the stress-strain states in the fiber, matrix and
their interface. In contrast, in the micromechanical approach, the constituents and their
interface can be definitely considered to predict the overall response of the composite as
well as the damage initiation and propagation in the composite. (Chen 2000)
Cheers,
Tim