This paper presents the preliminary results of an investigation into the use of a sandwich shell without spars or ribs as a horizontal tail on a light general aviation aircraft. The primary motivations for designing such a structure are simplicity, ease of construction, and the reduction of fasteners.
The structure investigated consisted of two symmetric, quasi-isotropic sandwich skins with four plies of graphite-epoxy unidirectional prepreg per face and 10-mm honeycomb core laid up in an airfoil-shaped mold and cured at 120°C. The two halves were joined by adhesive to form a tapered cantilevered beam with airfoil cross section truncated at the 56.7% chord line. Two internal ribs were installed at the fuselage contour near the center of the tail to provide attachment to a frame. Symmetric point loads were applied to produce varying amounts of torsion and bending, and the results were compared with analysis from a NASTRAN model using elements with composite material properties. The structure was loaded to failure, and the results were compared with analysis from a NASTRAN buckling routine.
A structurally intact portion of the failed test piece was modified to create one half of a fullsize forward-swept tail. Again, experimental results from various bending and torsion loads were compared with analysis from a comparable NASTRAN model.
Chordwise and spanwise strain plots indicated that the sandwich skins carry the bending loads, whereas the reinforced areas at the leading and trailing edges act as shear webs. Good correlation between the experimental strains and NASTRAN analysis was obtained.