As a start I am analyzing a fuselage structure. I realize that minimum gauge rules but I would like to calculate and get a feel for the margins involved.

I am considering a 0.5inch core (styrofoam 2lbs) with 2 and 3 layers of rutan uni and bid. I am using flat plate theory to calculate maximum loads. Then I am checking buckling of the sandwich panels- for this I assume the sides top and bottom are simply supported on the corners. Finally I check for wrinkling. in general I have found buckling to be critical. I am assuming the width of the panels to be 36inches and no intermediate bulkheads between tail and wing attach (is this typical?), i.e. long panels.

Assuming a layup of (0uni;45bid;0.5"foam;-45bid;0Uni) of rutan fabric and 2lbs/ft^3 styrofoam

D11 | 8,427.15 | Ec | 1,500.00 | |

D22 | 3,606.83 | Es | 1,000.00 | |

D12 | 1,651.11 | t | 0.54 | |

D66 | 1,990.12 | tf | 0.04 |

I get the following results:

N max -1100 lbs per inch (ultimate compression)

N wrk 359.1226411 lbs per inch (Wrinkling)

N bkl 169.741159 lbs per inch (Buckling)

Hence buckling is critical and I can take at most 359lbs/inch in compression before the top buckles. Is this in line with expectation? In my old case I have a 1000lbs load on the stabilizer and a 10 ft tail lever arm. This results in about ~100lbs per inch of compression load on the top panel so I have significant (~70%)margin. Reasonable? I know this is a little technical but any comments are appreciated.