Feature Article Figure 8. Panel compression test setup. Several strain gauges were positioned on different locations of the panel to monitor and record the local deformation caused by the compressive load. As anticipated, the strain recorded on the skin was considerably higher compared to the ribs, indicating the stiffness mismatch. The stain recorded on different ribs was very similar which confirmed the uniformity of loading of the panel. A closer look at the recorded strain values revealed that the highest deformation was recorded on the skin around the central part of the panel. For loads over 80 kN strain values from gauges located at the back of the skin started to decrease while the ones on the front continued to increase. This change was interpreted as local buckling of the skin towards the opposite side of the panel where the ribs were located. The low difference between strain values recorded from front and back strain gauges located close to the edge of the panel showed that the anti-buckling guides were preventing the panel from buckling significantly at its edges. The load on the fuselage panel increased until failure occurred at around 150 kN. A photograph of the failed panel is presented in Figure 9. The failure occurred at the top part of the panel, in the form of skin-rib debonding followed by skin damage. From the strain data recorded from the ribs it was evident that the strain gauges at the top part started to show some non linearity from a load level of 120 kN. As the structural failure of the panel occurred around the same location it can be postulated that the recorded nonlinearity is directly associated to the debonding failure of the skin and the ribs. 30 Level 4 Testing Finally, full scale testing on a scaled down prototype of the integrated lattice fuselage structure was performed at ambient conditions only. Finite element analysis was used to design the test fixture. The objective was to build a test fixture, which would have sufficient stiffness and be robust enough to apply the loads without being excessively heavy and complicated to manufacture and transport. Figure 10 shows the geometry and stress results of the final design of the test fixture at an applied load of 75 kNm. The fuselage section was subjected to bending in order to measure its deformation response, and use the collected data to validate the created numerical models. The load was applied via constant displacement (0.75 mm/min) of the hydraulic powered actuators located at each end of the test fixture. An overview of the load application setup is presented in Figure 11. The two actuators were synchronised and their load-displacement curves were recorded Figure 9. Failure of composite panel in compression. SAMPE Journal, Volume 52, No. 2, March/April 2016