a. An aircraft component is fabricated from an aluminium alloy that has a plane strain fracture toughness of 25 MPaym. It has been determined that fracture results at a stress

of 270MPa when the maximum (or critical) internal crack length is 1.5 mm. For this same component and alloy, will fracture occur at a stress level of 400 MPa when the maximum internal crack length is 1.0 mm? Why or why not? Note: the value of Y + 1. b. The fatigue data for a brass alloy given in table Q2b below.are i.Make an S-N plot using these data. ii.Determine the maximum stress level for which an infinite number of cycles will not cause failure. Demonstrate how this was obtained. ii.Determine the fatigue life at stress amplitude of 210 MPa and 160 MPa and explain the meaning of your solutions. Estimate fatigue strengths at 1.6 x 105 and 6 x 106 cycles and explain the implication of your answers. Demonstrate how this was obtained.iv. C. Steady-state creep rate data are given in Table Q2c below for nickel at 1000°C. If it is known that the activation energy for creep is 272,000 J/mol, compute the steady state creep rate at temperature of 1500°C and a stress level of 30 MPa. d. In table Q2d, weight gain-time data for the oxidation of copper at an elevated temperature ore tabulated i.Determine whether the oxidation kinetics obey a linear or parabolic rate expression. ii.Now compute W after a time of 800 min.