Design And Analysis Of Composite Stiffened Fuselage Panel

The main condition for ensuring high efficiency of the aviation technique is the intercalation of innovations in the field of design,technology and structural strength calculation.One of the most important areas in the modern aircraft industry is the use of polymer composite materials(PCM),which allows to create a lighter design and give the new properties for structure.Over the past 20 years,polymer composite materials have firmly occupied one of the main places among the structural materials in aviation.Due to the significant advantages in specific strength and stiffness,an especial combination of structural,thermal,technical properties,PCM are increasingly used in aircraft structures.The high potential for possible weight reduction of composite fuselage sections is due to the predominant use of carbon fiber,the development of optimal design and technological solutions,the development of technological processes corresponding to the characteristics of these materials and the specifics of the structures.Hence the relevance of this work,which are considering the design issues of the fuselage`s compartment of a wide-body long-haul aircraft.The process of designing a regular fuselage section was performed,the structure arrangement of this section was selected.Also,the analysis of the elements most susceptible to destruction and presents possible methods of monitoring the state of the structure.The purpose of this work is to find the best structural arrangement of the stiffened panel of the fuselage,strength calculation and determination of the geometric characteristics of the structural elements for a wide-body longhaul aircraft.To achieve this purpose,the design calculations of the polymer composite material fuselage section was made based on the mixture rule and specified laying of PCM,the required thickness of the skin from pressurization conditions in sealed compartment of the fuselage,the stringer set and frames,panel as a part of regular type of fuselage compartment was determined.During to calculation certain items of the FAR(federal aviation rules)were taken into account,which should provide the necessary reliability.More detail stress analysis of the lower panel of the fuselage was made taking into account the loads acting on the panel which was also obtained.Buckling analysis included compression buckling,shear buckling,bending buckling,crippling stress,Euler column buckling.As a result of this analysis the clear mock-up was made in Catia software.Verification based on comparison of results which was obtained from analytical calculation and calculation made with FEM made in Abaqus software proved accuracy of analytical research.