Question 1 of 8: Replacing 2024-T4 with 7075-T6 on a stretched business jet

Question 1: Replacing 2024-T4 with 7075-T6 on a stretched business jet (20 marks)

Question text not reproduced: the examination questions are © Engineers and Geoscientists BC. Open the official past paper (linked at the top of this page) to read the question, then follow the worked solution below.

Given. Lower wing skin panels, currently 2024-T4 (an Al–Cu–Mg alloy, solution treated and naturally aged), proposed to become 7075-T6 (an Al–Zn–Mg–Cu alloy, solution treated and artificially aged to peak strength). The extended version of the aeroplane weighs 40 per cent more than the original 12-passenger design. The panels see fluctuating tension in flight and fluctuating compression on the ground.

Find. The engineering case the recommending engineer would have made, argued from the physical and mechanical properties of the two alloys, together with an honest account of what is given up in exchange.

The load case first, because it decides everything else. A wing is a beam built into the fuselage and loaded by lift. In flight the upper surface is in compression and the lower surface is in tension; on the ground the wing hangs under its own weight and fuel load and the sense reverses, so the lower skin sees compression. A 40 per cent heavier aeroplane carries 40 per cent more lift at the same load factor, and because the design gross weight, the wing span and the taper are all bound together, the bending moment at the wing root scales roughly with the weight while the section modulus available in the existing skin does not. In round numbers the stress in an unchanged skin rises by about two fifths. Either the panel gets thicker, which returns much of the weight the stretch was trying to carry as payload, or the material gets stronger. The engineer chose the second route, and on static strength the case is strong: 7075-T6 has a 0.2 per cent proof stress of about 503 MPa against roughly 324 MPa for 2024-T4, and an ultimate strength of about 572 MPa against 469 MPa. That is a 55 per cent gain in yield, so the same limit load can be carried in a panel roughly a third thinner, and the compressive yield strength — which is what resists the ground-load buckling of a lower skin between stringers — improves in the same proportion.

The physical properties are almost interchangeable, which is what makes the substitution attractive. The densities are 2.78 and 2.81 g/cm³, so the swap costs about one per cent in specific weight and returns far more than that in allowable stress. Young's modulus is 73.1 GPa for 2024 and 71.7 GPa for 7075, a difference of two per cent, so any part of the structure sized by stiffness rather than strength — skin panel flutter, aileron reversal, the deflection that sets fuel-tank sealing — is essentially unaffected, and no aeroelastic requalification is forced by the change. Thermal expansion, thermal conductivity and electrical conductivity are close enough that the fastening, bonding and lightning-protection schemes carry over. Both alloys are clad in practice (Alclad 2024 and Alclad 7075) with a thin layer of pure or 7072 aluminium for corrosion protection, and both are joined by riveting rather than welding, so the manufacturing route is unchanged. In short, the engineer is proposing to buy strength with almost no penalty in weight, stiffness or process.

What is given up: damage tolerance. The 7xxx alloys owe their strength to a very fine, dense dispersion of MgZn₂ (η') precipitates formed on artificial ageing, and that same microstructure gives them poor resistance to crack growth. Plane-strain fracture toughness in the longitudinal-transverse orientation is about 37 MPa·m^0.5 for 2024-T3/T4 sheet against roughly 26 MPa·m^0.5 for 7075-T6, and the fatigue-crack growth rate at a given stress-intensity range is appreciably higher. For a tension-dominated, fatigue-critical location this is exactly the wrong trade. Since $K = Y\sigma\sqrt{\pi a}$, the critical crack length at a given stress scales as $\left(K_{Ic}/\sigma\right)^{2}$, so raising the working stress by two fifths and dropping the toughness by a third shrinks the tolerable crack by roughly a factor of four — (26/37)² × (1/1.40)² = 0.25. In a damage-tolerance regime that translates directly into shorter inspection intervals over the life of the airframe. This is precisely why traditional practice puts 2024 on the lower (tension) surface and reserves 7075 for the upper (compression) surface, where crack growth is retarded by the compressive part of the cycle and the driving requirement is compressive strength and stability.

Three further trade-offs deserve to be on the record. First, stress-corrosion cracking: 7075 in the peak-aged T6 condition is notoriously susceptible in the short-transverse direction, and any part where a sustained tensile stress acts through the thickness — a machined step, an interference-fit fastener, a shimmed joint — is at risk. The standard mitigation is to specify an over-aged temper such as T73 or T7351, which trades ten to fifteen per cent of the strength for a large gain in stress-corrosion and exfoliation resistance and some recovery of toughness; T76 sits between the two. Second, formability: 2024 in the freshly quenched W condition can be formed and then allowed to age naturally, whereas 7075-T6 is hard, has an elongation of about 11 per cent against 19 per cent for 2024-T4, and must be formed in the annealed or W temper and heat treated afterwards, with the distortion and residual stress that implies for a large curved skin. Third, cost and supply: 7075 plate and sheet carry a modest premium and, in the wing-skin thicknesses of interest, longer lead times.

Verdict. The recommendation is defensible and its motive is transparent: a 40 per cent heavier aeroplane needs about 40 per cent more strength from the same skin thickness, and 7075-T6 supplies it at essentially no weight, stiffness or manufacturing penalty. It is nevertheless only half an argument, because it optimises for static strength in a location whose real design driver is fatigue and damage tolerance. The professional answer is to accept the alloy change but not the temper: specify 7075-T73 or T7351, or better still a modern damage-tolerant alternative such as 7150-T77 or 2524-T3, run a fresh crack-growth and residual-strength analysis for the new load spectrum, and set the inspection intervals from that analysis rather than from the old ones.