Aerospace Materials & Composites

Choosing materials where every gram counts. Learn specific strength and stiffness, aluminium alloys (2xxx, 7xxx, Al-Li), titanium, steels, nickel superalloys, ceramics and CMCs, polymer-matrix composites (carbon/glass/aramid), lamina properties and the rule of mixtures, laminate lay-ups, manufacturing (prepreg, autoclave, AFP, RTM), damage and inspection, and material selection — with Python comparing metals and computing CFRP ply properties.

Aircraft materials have changed from wood and fabric, to aluminium (1930s onward), to carbon-fibre composites — now about 50% of the structural weight of the Boeing 787 and Airbus A350. Engines rely on titanium and nickel superalloys; spacecraft use everything from aluminium–lithium tanks to ceramic heat shields.

Designers compare materials by properties divided by density (specific properties), because what matters is how light a part can be for a required stiffness or strength.

An analogy

A composite is like reinforced concrete at microscopic scale: strong, stiff fibres (like steel rebar) carry the load in their direction, and a softer matrix (like concrete) holds them in place and transfers load between them.

Metals in aircraft

The main alloy families:

Composites

Fibre-reinforced polymers combine fibres (carbon — stiff and strong; glass — cheap and tough; aramid/Kevlar — impact-resistant) with a matrix (epoxy thermosets, or thermoplastics like PEEK). A single ply is highly anisotropic; laminates stack plies in chosen directions to tailor stiffness — including aeroelastic tailoring of wings. Sandwich panels (skins on honeycomb or foam cores) give very high bending stiffness per weight.

Manufacturing

Prepreg plies are laid up by hand or by automated fibre placement (AFP) / tape laying (ATL), vacuum bagged and cured in an autoclave (~180 °C, 6–7 bar). Out-of-autoclave processes (RTM, infusion) reduce cost. Parts are drilled and fastened or co-cured/co-bonded to reduce fasteners. Metals are machined from thick plate (buy-to-fly ratios up to 10:1), forged, or increasingly additively manufactured.

Damage, inspection and selection

Composites don't dent visibly: barely visible impact damage (BVID) can cause internal delaminations that reduce compression strength by 40–60%, so designs are certified with damage present. Inspection uses ultrasonic C-scan, thermography and X-ray CT. Lightning strike protection (copper mesh), moisture absorption, hot-wet properties and repairability are key concerns. Ashby charts and merit indices (E/ρ, E^(1/2)/ρ, E^(1/3)/ρ) guide selection systematically.

Formulas

Worked example: Rule of mixtures

Given: E_f 230 GPa, E_m 3.5 GPa, V_f 0.6

Find: E₁.

Answer: 139.4 GPa

Worked example: Weight saving

Given: Stiffness-critical tie replaced Al → UD CFRP

Find: Mass ratio.

Answer: ≈ 0.29 → 71% lighter

Worked example: Panel merit index

Given: Al: E 72 GPa, ρ 2810; CFRP quasi-iso: E 50 GPa, ρ 1560

Find: E^(1/3)/ρ ratio.

Answer: CFRP ≈ 1.6× better for buckling-critical panels

Specific properties and a carbon/epoxy ply

# Specific stiffness/strength and rule of mixtures for a UD carbon/epoxy ply
mats = {"Al 7075-T6": (71.7e9, 503e6, 2810), "Ti-6Al-4V": (114e9, 880e6, 4430),
        "Steel 4340": (205e9, 1470e6, 7850)}
for n, (E, s, rho) in mats.items():
    print(f"{n:11s}: E/rho = {E/rho/1e6:5.1f} MN*m/kg, sigma/rho = {s/rho/1e3:6.1f} kN*m/kg")
Ef, Em, Vf = 230e9, 3.5e9, 0.6
E1 = Ef * Vf + Em * (1 - Vf)
E2 = 1 / (Vf / Ef + (1 - Vf) / Em)
rho = 1800 * Vf + 1200 * (1 - Vf)
print(f"UD CFRP: E1 = {E1/1e9:.1f} GPa, E2 = {E2/1e9:.2f} GPa, rho = {rho:.0f} kg/m3, E1/rho = {E1/rho/1e6:.1f} MN*m/kg")
# Al 7075-T6 : E/rho =  25.5 MN*m/kg, sigma/rho =  179.0 kN*m/kg
# Ti-6Al-4V  : E/rho =  25.7 MN*m/kg, sigma/rho =  198.6 kN*m/kg
# Steel 4340 : E/rho =  26.1 MN*m/kg, sigma/rho =  187.3 kN*m/kg
# UD CFRP: E1 = 139.4 GPa, E2 = 8.55 GPa, rho = 1560 kg/m3, E1/rho = 89.4 MN*m/kg

A striking fact: almost all structural metals have the same specific stiffness (~26 MN·m/kg). Unidirectional carbon fibre is ~3.5× better along the fibres — but only 8.6 GPa across them, which is why laminates stack plies at 0°, ±45° and 90° (quasi-isotropic laminates reach ~50 GPa).

Common mistakes

Applications

Key takeaways

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