The rib-skin construction using carbon fiber composites has been an attractive design option for aircraft control surfaces like aileron, elevator, rudder, flap etc. The major concern in such structures is the debonding between skin and rib flange which can occur due to low velocity impacts like tool drops, run way debris etc. Such debonds which occur in service are barely visible and may not get detected till the next inspection schedule. The integrity of the structureudin the intervening period is of great concern to the designers. In the present work, the structural integrity of a composite flap structure having multiple debonds at the rib skin interfaces under fatigue loading has been addressed. The flap is subjected to cyclic loading atuddesign limit load for 110000 cycles using a whiffle tree mechanism. The strains have been monitored at different locations to understand the behaviour of structure during the test using strain gauges and Fiber Bragg Grating (FBG) sensors. Ultrasonic A-scan was used to monitor the defect growth after each block of 1000 cycles. The growth of debonds was not significant during the fatigue testing. The strain levels did not change appreciably throughout the testudperiod indicating the damage tolerance capacity of the flap structure. The low growth of debonds was attributed to the low level of strains in the structure since the flap design is driven by stiffness considerations.
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