The transverse-loading response of laminated composite shell structures is studied experimentally and numerically. Monolithic graphite/epoxy shell structures having layups of [§ 45n/0n]s (n= 1;2;and3) closely represent commercial fuselage structures in both geometry and boundary conditions. A combined experimental and numerical approach is used to assess shell response to centered transverse loading. Experimentally, load-dee ection response and mode-shape evolutions are measured and damage resistance characterized via dye-penetrant enhanced x radiography and sectioning. Nonlinear e nite element analyses including buckling and dynamic collapse are conducted for comparison to the experimental data. Modeling results allow a more ree ned interpretation of observed bifurcation phenomena, particularly premature transition to a secondary equilibrium path attributed to geometric imperfections. A novel e nite element technique introduced in previous work is found to be superior to traditional methods for identifying and traversing bifurcation points in this work. A simply supported axial boundary conditionisfound to givea much morecomplex buckling response (bifurcationandlimit-pointbuckling, as well as dynamic collapse ) than specimens with a free axial edge (bifurcation or limit-point buckling ). Experimental and numerical comparisons for the range of thicknesses considered indicate that elasticity of the in-plane boundary condition and transverse shear effects need further consideration. Observed shell damage is typical of that observed for composite plates. Results of this work give new insight into the response of composite fuselage panels to damaging transverse events, particularly in regard to instability/buckling behavior.