Design Method Of Composite Laminate Structure And Research On The Strength Of Scarf Patch Repair

In recent years, the use of composite materials in large aircraft structure increasedrapidly for their high specific strength, modulus and outstanding designability characte-ristic. Study of composite material structure design, structure strength check and therepair of damage structures have become important works in aircraft structure designand routine maintenance. Focusing on the problem of structure design of compositematerial I-beam, strength check and composite laminated plates scarf patch repair, thefollowing research work is elaborated in this paper:(1) The research trend of composite material structure design and scarf patch repairis presented. The basic principles of composite structural mechanics, the classicallaminated plate theory, the laminated plate strength theory and the basic principle offinite element method (FEM) are introduced.(2) The layer structure of an I-beam is designed. The internal force distributions offlange and web under specific external load are calculated according to the I-beammechanical characteristics. Stacking-sequence of flange and web is determined based onthe composite material stacking principle. The composite ply direction and quantity areiteratively calculated using MATLAB software according to the classical laminatedplate theory and the maximum stress failure criterion.(3) The effect of extreme load on different layup methods of I-beam is researchedunder the condition of the same ply orientation and quantity, and the optimal layup isobtained. Relative errors of ultimate load between experimental values and the results ofsimulation are calculated respectively using FEM. The flange and web of I-beam usinginner C layup method process has the best mechanical performance, and its relativeerror is8.93%.(4) Static loading experiment of an I-beam with inner C layup method is carriedout. The strain values in two typical location of upper and lower flange are collectedusing strain gages. The results of simulation and the experimental values are compared.The relative errors of strain are calculated. The relative error is31.76%for upper flangeand12.66%for lower flange. Lower flange is under tensile loading, which leads tomore accurate experimental results. Compression of upper flange causes local buckling,which leads to larger relative error.(5) Five different finite element models of scarf patch repair laminate are establish- ed, and the stress calculation and strength recovery analysis are applied. The lowestcoefficient of strength recovery occurs in the laminate plate with penetrating damagerepaired by scarf patch method. The laminate is weakened by the penetrating damage.The highest coefficient of strength recovery occurs in the laminate with non-penetratingdamage repaired by scarf patch method with two additional plies, even greater than theextreme strength of the original undamaged laminate. Considering that the patchadhesive bonder has certain influence to the model, extreme strength of the repairedmodel will be increased.(6) Wet molding process method of composite material is researched. An I-beamand the composite structure of a small unmanned aircraft is designed and manufactured.Control systems and power plant are assembled and debugged. The first flight of the fullcomposite structure unmanned aircraft is carried out successfully.