The newly developed airplanes are using composite laminates to replace the metal alloys for different components, such as the fuselage and the wings. The major advantage of the composite materials is to reduce structural weight which results in reducing the fuel consumption. The aim of this project is to investigate the structural integrity of an airplane fuselage, which uses various types of carbon composite laminates under the static loading of the cabin pressurization. The research is performed using the finite element method and the HYPERMESH commercial software with a composite tool to change the thickness and the orientation of carbon fiber laminates used in the facesheet of the sandwich structure. Three different orientations/stacking sequence of the HexPly 8552 AS4 carbon fibers with two honeycomb cores: Hexagonal Al and Nomex. The results show that the composite material using the HexPly 8552 carbon fiber oriented at angle 30 and angle 45 and the Nomex Honeycomb core of a total laminate thickness of 15.875mm outperform all other thicknesses and orientations in regards to the static loading failure.