A wing panel of a supersonic aircraft is made from a titanium alloy that has yield strength 1035 MPa and toughness KIc = 55 MPa√m. It is 3.0 mm thick, 2.40 m long, and 2.40 m wide. In service it is subjected to a cyclic stress of ±700 MPa, which is not enough to cause yielding but does cause gradual crack growth of a preexisting crack normal to the loading direction at the edge of the panel. Assume that the crack is initially 0.5 mm long and grows at a constant rate of da/dN = 120.