Damage Repair Of The Glass Fiber Reinforced Aluminum Laminates And The Performance After Damage Repair

The development and manufacturing of large aircraft brings forward the impending needs for high-strength and lightweight advanced composite materials.Glare(Glass Reinforced Aluminum laminates)has been treated as the crucial structural materials for fuselage and wing skin owing to its excellent fatigue,good damage tolerance as well as the outstanding contributions to weight loss.However,under various kinds of loading,aircraft body structure will inevitably form corrosion damage,impact damage and fatigue damage.If not repaired timely,can cause structural damage,affect flight safety.In this paper,three kinds of damage modes,repair process and mechanical properties of the laminated plate are studied.First of all,according to the preparation process of fiber metal laminates,we get GLARE4 A laminate;Secondly,damage specimen is prepared according to the damage form;Then,the process of cementing repair was developed to determine the material,size and thickness of the patch and the curing process of the adhesive.Finally through the static tensile test,drop weight impact test and fatigue crack growth rate test to compare the structural response of the specimens before and after repair,at the same time reveals the repair after the damage mechanism in the static tensile test specimens.The following conclusions are drawn:The GLARE laminates' bonding repair process mainly includes: repair methods,choice of adhesive,curing process and the preparation of the patch.The bearing capacity of bonding repair are higher than mixed with the glue riveting repair and riveting repair.Choose J-150 adhesive,of which curing process using 80? ×4h + 120? ×2h and the curing mode is cocuring.The patch shape is square,of which thickness is 1.8 mm and the size of the patch is 3 times of stress damage area;The static tensile test of the bonding repair shows that the energy of the damage to the specimen retention rate is only 37.21%,the parameters shows that damage causing the deterioration of tensile load and deformation capacity at the same time.After bonding repair,energy retention rate of the specimens mentioned to 59.99%,the energy recovery rate was 22.77%;Bonding repair after the failure mode of damage in the static tensile test specimens are: overall debonding patch;Crack on the earliest produced in damage hole;After the repair layer board of load displacement curve,sudden drop in load reaches 14920.2 N,then the curve continue to rise until failure;Due to the stress concentration on damage hole initiation crack first,with the increase of load,rubber shear deformation damage and failure and the patch can't reflect the enhancement function of load by the motherboard bearing all to final failure;During the whole experiment,there was no damage to the patch,and the internal structure was intact,indicating that the strength of the patch was sufficient,the patch was unglued,and there was no defect in the patch,and the structural integrity was maintained;The energy of hammer impact is 5J,15 J and 25 J respectively,corresponding to three types of damage.Research results show that the deformation of specimen after repair center less than not repaired specimens,the displacement time history curve of the specimens after repair will appear in the sawtooth platform,this is because the specimen within the fiber damage caused by impact load fluctuate;Due to the increase of stiffness,the peak of impact load of the post-repair specimens in the load time curve is higher than that of unrepaired specimens,however during the 25 J energy shock,the specimen was punctured directly,and the back of the repaired specimen was cracked by the stretching action,but it was not penetrated.The length of the prefabricated crack was 10 mm in the test center,simulating fatigue damage,and compared the fatigue crack propagation rate experiment before and after repair to evaluate the quality of the repair.a-N graph is obtained by using the cyclic frequency N as the horizontal coordinate and the crack length a as the longitudinal coordinate system.The value of da/dN is obtained by using origin,i.e.the rate of fatigue crack propagation.It can be seen that the fatigue crack propagation rate of the test piece before and after the repair is relatively stable.The fatigue crack expansion rate of the repaired specimen is 1.48E-5,which is lower than the da/dN value of the unrepaired test piece.In the stress intensity factor ?K as the abscissa,fatigue crack growth rate da/dN as the ordinate double logarithmic coordinate system,using origin software fitting curve C and m values,can be seen from the curve in the crack initiation and crack extension stage,repair the specimens after fatigue crack propagation rate is below not repaired specimens.