Aerospace materials and composites

In short

Aerospace materials engineering is the selection and processing of metals, polymers, ceramics and fibre-reinforced composites so that a structure carries flight loads at the least possible weight, for the required life and temperature. The driving figures of merit are not strength and stiffness alone but specific strength (σ/ρ) and specific stiffness (E/ρ), together with fatigue, fracture toughness, corrosion and temperature capability. Composites — fibres in a matrix — now form 50% or more of the structural weight of modern airliners (Boeing 787, Airbus A350) and a growing share of a launch vehicle's motor cases plus most of its payload fairings and heat shields. The topic carries heavy weight in GATE AE (materials, laminate micromechanics) and in technical rounds at HAL, ISRO/VSSC, DRDO (ADA, DMRL, ADE), NAL, Tata Advanced Systems, Godrej Aerospace, L&T Precision and Safran/Collins/Boeing–Airbus India centres. On the job it appears as material selection, process-sheet reading, NDT dispositioning and repair engineering.

Explained

Why specific properties decide aerospace selection

• Aircraft mass drives fuel burn; a kilogram saved on a structure saves many more over the life of the vehicle, so the metric is property per unit density, not property alone.
• Specific strength = σ_ult/ρ (N·m/kg); specific stiffness = E/ρ. Steel and aluminium have almost the same E/ρ (~26 MN·m/kg), so steel buys no stiffness advantage per kilogram — only strength and temperature.
• Buckling-critical parts (skins, webs) are selected on E^1/2/ρ or E^1/3/ρ (column versus plate-bending index, depending on whether thickness or geometry is free); tension-critical parts on σ/ρ. This is why thin aluminium skins with stringers beat a thick steel skin.
• Other mandatory screens: fatigue strength and crack-growth rate (aircraft are fatigue-critical), fracture toughness K_IC (damage tolerance), corrosion resistance, creep and oxidation above ~0.4 T_melt, cost and manufacturability, and availability/ITAR-free sourcing — a real constraint for Indian programmes.
• Damage tolerance philosophy (today's norm, per CAR/FAR 25.571 and MIL-A-87221 practice): assume a crack exists, size the structure so it will not grow to failure between inspections. Earlier philosophies were *safe-life* (retire at a fatigue life) and *fail-safe* (alternate load path).

Metallic aerospace materials

• Aluminium alloys — the airframe workhorse; ρ ≈ 2700–2800 kg/m³, E ≈ 70–73 GPa.
  • 2xxx (Al–Cu, e.g. 2024-T3/T351): σ_y ≈ 325 MPa, σ_u ≈ 470 MPa, excellent fatigue and toughness — fuselage skins, lower wing skins (tension-dominated).
  • 7xxx (Al–Zn–Mg–Cu, e.g. 7075-T6): σ_y ≈ 500 MPa, σ_u ≈ 570 MPa, higher strength but lower toughness and stress-corrosion prone in T6 — upper wing skins, spars, fittings (compression-dominated). T73/T7351 over-ageing trades ~10% strength for stress-corrosion-cracking resistance.
  • Al–Li (2195, 8090, 2050): ~3% lighter, ~10% stiffer; used in cryogenic tanks (Space Shuttle/SLS external tank, launch-vehicle tanks).
  • Clad sheet (Alclad) — pure-Al layer rolled on for corrosion protection; anodising and chromate-free primers for surface protection.
  • Strengthened by precipitation (age) hardening: solution treat → quench → natural (T3/T4) or artificial (T6) ageing; GP zones and θ″/η′ precipitates block dislocations.
• Titanium alloys — ρ ≈ 4430 kg/m³, E ≈ 110 GPa, usable to ~400–500 °C, superb corrosion resistance and galvanically compatible with carbon fibre.
  • Ti-6Al-4V (α+β, Grade 5): σ_y ≈ 830–900 MPa — engine fan blades and discs, landing-gear parts, wing-box fittings, and almost all metal fittings bolted to CFRP (aluminium would corrode).
  • β alloys (Ti-10V-2Fe-3Al) for high-strength forgings; α alloys for weldability and creep.
  • Costly and difficult to machine (low thermal conductivity, chemical reactivity); needs inert-atmosphere welding. India's DMRL and Mishra Dhatu Nigam (MIDHANI), Hyderabad supply aerospace titanium and superalloys.
• Steels — used only where strength, stiffness or wear demand it: 300M / 4340 / AISI 4130 for landing gear and engine mounts (σ_u up to ~1900 MPa), maraging steels (M250/M300) for solid rocket motor cases (ISRO's S200, PSLV stages — made at MIDHANI/VSSC), precipitation-hardening stainless 17-4PH, 15-5PH for fittings.
• Nickel-base superalloys — Inconel 718, Nimonic 90, single-crystal CMSX/DS alloys for turbine blades; γ′ (Ni₃Al, Ti) precipitation strengthening, usable to ~1000–1100 °C with thermal barrier coatings (YSZ) and film cooling. Directional solidification removes transverse grain boundaries that fail by creep.
• Magnesium alloys (ρ ≈ 1800 kg/m³) — gearbox and accessory casings; limited by corrosion and flammability rules.
• Ceramics and C–C: carbon–carbon and SiC/SiC for nose caps, brake discs, nozzle throats; ablative silica-phenolic and carbon-phenolic for re-entry and rocket nozzles (ISRO's Gaganyaan crew-module heat shield uses an ablative tile system developed with VSSC).

Composites: constituents and classification

• A composite is a macroscopic combination of two or more chemically distinct phases with a recognisable interface, where the reinforcement (high strength/stiffness) carries load and the matrix binds, transfers load through shear, protects fibres and sets temperature and environmental limits.
• Classification by matrix: PMC (polymer — the bulk of airframe use), MMC (metal matrix, e.g. Al–SiC for struts, Ti–SiC fan blades), CMC (ceramic matrix, SiC/SiC turbine shrouds), CCC (carbon–carbon).
• By reinforcement form: continuous (unidirectional, woven fabric, braid), short/chopped fibre, particulate, whisker, and structural (laminate, sandwich).
• Fibres:
  • Carbon: standard modulus T300/T700 (E ≈ 230 GPa, σ ≈ 3500–4900 MPa), intermediate IM7 (E ≈ 276 GPa), high modulus M55J/M60J (E > 500 GPa, lower strain). PAN-based for structures, pitch-based for stiffness and thermal paths.
  • Glass: E-glass (E ≈ 72 GPa, cheap, electrically insulating — radomes, fairings, GA aircraft), S-glass (E ≈ 86 GPa, higher strength).
  • Aramid (Kevlar 49): E ≈ 131 GPa, low density 1440 kg/m³, superb impact and energy absorption — armour, ballistic panels, fuel-cell shells; weak in compression and absorbs moisture.
  • Boron, SiC, alumina, UHMWPE (Dyneema) for special cases.
• Matrices: epoxy (120–180 °C cure, best adhesion and properties — 180 °C cure prepregs like 977-3, 3501-6), BMI (to ~230 °C), polyimide/PMR-15 (to ~300 °C), phenolic (fire, smoke, toxicity compliant — cabin interiors per FAR/CAR 25.853), thermoplastics PEEK/PEKK/PPS (tough, weldable, infinite shelf life, used in A350 clips and Gaganyaan-class hardware).
• Interface/sizing governs load transfer; poor interfacial bond gives low transverse and shear strength even with excellent fibres.
• Anisotropy and orthotropy: a unidirectional ply is far stronger along the fibre (direction 1) than transverse (2). Hence engineered stacking: 0° for axial load, ±45° for shear and torsion, 90° for hoop/transverse, and a balanced, symmetric layup to avoid coupling and warpage.

Manufacturing processes for composites

1. Hand lay-up / wet lay-up: fabric wetted with resin in a mould; cheapest, lowest fibre volume (~0.35–0.45), high voids — gliders, fairings, repairs.
2. Prepreg + autoclave: pre-impregnated unidirectional tape/fabric laid up, vacuum bagged (bleeder, breather, release film), cured under 6–7 bar and 120–180 °C. Gives V_f ≈ 0.55–0.65 and <1–2% voids — the primary-structure standard (HAL's Composite Manufacturing Division, Bengaluru; Tata Advanced Systems for Apache fuselage and C-295 work).
3. Resin transfer moulding (RTM) / VARTM / resin infusion: dry preform in a closed mould, resin drawn or injected; out-of-autoclave, good for large thick parts.
4. Filament winding: continuous tow wound on a rotating mandrel at a chosen helix angle — pressure vessels, rocket motor cases, COPVs. Netting analysis sets the optimum helix of ~54.7° for an isotropic-pressure cylinder.
5. Pultrusion: constant-section profiles pulled through a heated die — spars, stringers, floor beams.
6. ATL/AFP (automated tape/fibre placement): robotic heads lay tape on complex contours; the 787 barrel and A350 spars are made this way.
7. Compression moulding / SMC, BMC and injection moulding of short-fibre thermoplastics for brackets and secondary parts.
8. Sandwich fabrication: Nomex or aluminium honeycomb or foam core co-cured/adhesively bonded (FM-73, redux film) between thin facesheets; huge increase in bending stiffness per unit weight — floor panels, control surfaces, radomes.
9. Curing needs a correct ramp, dwell and cool-down; under-cure leaves low T_g, over-cure embrittles. Prepregs have an out-time and freezer shelf life that must be logged.

Failure, damage and inspection

• Micro-mechanical failure modes: fibre breakage, matrix cracking, fibre–matrix debonding, fibre pull-out, fibre microbuckling/kinking in compression, delamination (interlaminar separation — the dominant and most dangerous mode, driven by weak through-thickness properties).
• BVID (Barely Visible Impact Damage): a tool drop or hail can create internal delamination with no visible surface mark; compression-after-impact strength can fall 40–60%. Design allowables are therefore set by compression after impact (CAI), not pristine strength.
• Failure theories examined: maximum stress, maximum strain, Tsai–Hill and Tsai–Wu (interactive, quadratic), with first-ply failure then progressive-failure analysis.
• Environmental effects: moisture absorption plasticises epoxy and lowers T_g and hot/wet compression strength — hence the "hot/wet" design condition; UV degradation; galvanic corrosion when carbon contacts aluminium (needs a glass-fabric isolation ply and sealant).
• NDT methods: visual and tap test; ultrasonic A/C-scan and phased array (standard for delamination and porosity); thermography; X-ray/CT for honeycomb and water ingress; shearography; acoustic emission for proof tests. Coin-tap for disbond in sandwich.
• Repairs: scarf or stepped bonded patch repair for primary structure, bolted doubler where bonding is not qualified, resin injection for small disbonds — all per the SRM (Structural Repair Manual) and requiring DGCA-approved process and personnel.
• Joints: composites prefer bonded joints (no stress-raising holes) but certification often demands bolts ("chicken rivets"); bolted composite joints need large bearing area, titanium fasteners, and checks for bearing, shear-out, net-tension and pull-through.

Indian standards, regulation and practice

• Civil airworthiness in India is regulated by DGCA through the Civil Aviation Requirements (CAR); CAR Section 2 Series mirrors FAR/CS-25 requirements including structural strength, fatigue and damage tolerance (25.571) and cabin flammability (25.853). Military airworthiness is certified by CEMILAC with inspection by DGAQA; ADA designs and HAL builds the LCA Tejas.
• Material qualification follows MIL-HDBK-5/MMPDS (metals) and MIL-HDBK-17/CMH-17 (composites) for design allowables, with ASTM test standards (D3039 tension, D3410/D6641 compression, D2344 short-beam shear, D5528 Mode-I G_IC, D7136/D7137 impact and CAI, D2584 resin burn-off, D792 density).
• BIS publishes Indian Standards for aluminium wrought products (IS 733 for bars/rods/sections, IS 737 for sheet and strip) and for glass-fibre reinforced plastics; aerospace procurement, however, generally calls out AMS/MIL/Airbus-Boeing specifications rather than IS. Quote an IS number only from the current catalogue — several older aerospace-related IS have been withdrawn or superseded.
• Quality systems: AS9100D (aerospace QMS) and NADCAP accreditation for special processes (heat treat, NDT, composites, chemical processing) are commercial prerequisites for Indian Tier-1/Tier-2 suppliers.
• Workplace health and safety for resins, solvents and dusts now falls under the Occupational Safety, Health and Working Conditions Code, 2020, which came into force with the four Labour Codes on 21 November 2025, replacing the Factories Act, 1948 and the BOCW Act; hazardous-chemical handling additionally sits under the Environment (Protection) Act rules.
• Indian ecosystem: VSSC/ISRO (motor cases, nozzles, ablatives, composite payload fairings), DMRL and MIDHANI (titanium, superalloys, maraging steel), NAL (composite wing, Saras, autoclaves), HAL CMD, Tata Advanced Systems, Kineco Kaman, Godrej Aerospace, Bharat Forge, L&T.

Key points

• Specific strength / stiffness - σ/ρ and E/ρ; the true aerospace selection metrics, not absolute strength.
• 2024-T3 vs 7075-T6 - 2024 for fatigue/tension (fuselage, lower wing); 7075 for strength/compression (upper wing, spars).
• Age hardening - solution treat, quench, age; fine precipitates pin dislocations and give Al its strength.
• Ti-6Al-4V - the α+β workhorse; mandatory at CFRP interfaces because it is galvanically compatible with carbon.
• Superalloy creep control - γ′ precipitates plus directional solidification or single crystal; TBC and film cooling push gas temperature above the metal's melting point.
• Maraging steel - ultra-high-strength, low-carbon Ni steel for solid rocket motor cases (ISRO/MIDHANI).
• Reinforcement vs matrix - fibres carry load; matrix transfers shear, protects fibres and fixes the service temperature.
• Rule of mixtures - E₁ = E_f V_f + E_m V_m (Voigt, in-fibre); 1/E₂ = V_f/E_f + V_m/E_m (Reuss, transverse).
• Fibre volume fraction - ~0.6 for autoclaved prepreg, ~0.4 for hand lay-up; void content must stay under about 2%.
• Balanced and symmetric layup - prevents extension–shear and bending–stretching coupling (B matrix = 0) and cure warpage.
• Quasi-isotropic - [0/±45/90] gives in-plane properties independent of direction, roughly one-third of UD stiffness.
• Delamination and BVID - low interlaminar strength makes impact damage, often invisible, the governing design case (CAI allowables).
• Hot/wet condition - moisture lowers T_g and compression strength; the certification-critical environment.
• Sandwich construction - thin stiff facesheets on a light core give the highest bending stiffness per unit mass.
• Autoclave prepreg - 6–7 bar, 120–180 °C; the benchmark for primary structure; RTM/infusion is the out-of-autoclave route.
• Tsai–Hill / Tsai–Wu - interactive failure criteria for an orthotropic ply; max-stress/max-strain are non-interactive.
• Damage tolerance - assume a crack, size for inspection intervals; CAR/FAR 25.571 and CMH-17 allowables.
• Regulators - DGCA (CAR) for civil; CEMILAC certifies and DGAQA inspects military; AS9100D and NADCAP for suppliers.

Formulas and examples

Example 1 — Specific properties: which material for a tension-critical tie rod?

• Steel, titanium and aluminium sit within ±3% of one another on E/ρ — so a metal swap buys no stiffness per kilogram.
• Unidirectional carbon/epoxy gives ~3.5× the E/ρ and ~5× the σ_u/ρ of all three metals: hence the composite wing box.
• Check: 570×10⁶/2800 = 203.6×10³ N·m/kg = 204 kN·m/kg. 1800×10⁶/1580 = 1139 kN·m/kg. ✔

Example 2 — Rule of mixtures for a carbon/epoxy lamina. Given E_f = 230 GPa, ρ_f = 1800 kg/m³, E_m = 3.4 GPa, ρ_m = 1200 kg/m³, V_f = 0.60, no voids.

1. V_m = 1 − 0.60 = 0.40.
2. E₁ = 230(0.60) + 3.4(0.40) = 138 + 1.36 = 139.4 GPa.
3. 1/E₂ = 0.60/230 + 0.40/3.4 = 0.002609 + 0.117647 = 0.120256 → E₂ = 8.32 GPa (≈ 1/17 of E₁ — the anisotropy).
4. ρ_c = 1800(0.60) + 1200(0.40) = 1080 + 480 = 1560 kg/m³.
5. Fibre weight fraction W_f = 1080/1560 = 0.692 (69.2% by weight for 60% by volume).
6. Specific stiffness E₁/ρ_c = 139.4×10⁹/1560 = 89.4 MN·m/kg against 25.4 for 7075-T6.
7. Load share at equal strain: P_f/P_m = 138/1.36 = 101, i.e. the fibres carry about 99% of the longitudinal load.

Example 3 — Mass saving on a CFRP wing-skin panel. A panel must carry a tensile load of 1.2 MN over a 1.0 m width with a factor of safety of 1.5, so design stress must satisfy σ_design = σ_allow.

• Aluminium 2024-T3, σ_allow = 470/1.5 = 313 MPa → required area A = 1.2×10⁶/313×10⁶ = 3.83×10⁻³ m² → t = 3.83 mm. Mass per metre length = 3.83×10⁻³ × 1.0 × 2780 = 10.66 kg/m.
• CFRP (quasi-isotropic, σ_allow in load direction taken as 600/1.5 = 400 MPa after knockdown for impact and hot/wet) → A = 1.2×10⁶/400×10⁶ = 3.0×10⁻³ m² → t = 3.0 mm. Mass = 3.0×10⁻³ × 1580 = 4.74 kg/m.
• Saving = (10.66 − 4.74)/10.66 = 55.5%. Note the composite value already carries a large knockdown — pristine UD strength would be ~1800 MPa, and using that would be the classic exam error.

Example 4 — Filament-wound COPV / rocket motor case. A cylindrical case of internal radius R = 0.5 m must hold p = 6 MPa; carbon/epoxy allowable σ_allow = 1000 MPa (with safety factor already applied).

1. σ_hoop = pR/t = 6×0.5/t = 3/t (MPa with t in m); σ_long = pR/2t = 1.5/t.
2. Hoop governs: t = 3/1000 m → t = 3.0 mm.
3. Optimum helix angle: tan²α = σ_hoop/σ_long = 2 → α = arctan(1.414) = 54.74° to the axis (the "geodesic" or netting-analysis angle).
4. Shell mass (ρ = 1580 kg/m³, length 2 m): m ≈ 2πRtρL = 2π(0.5)(0.003)(1580)(2) = 29.8 kg. An equivalent maraging-steel case (σ_allow 1300 MPa, ρ 8000 kg/m³, t = 3/1300 = 2.31 mm) would weigh 2π(0.5)(0.00231)(8000)(2) = 116 kg — the reason ISRO moved from metallic to composite motor cases and COPVs.

Example 5 — Critical crack size and inspection (damage tolerance). A 7075-T6 fitting has K_IC = 24 MPa√m and operates at σ = 200 MPa. Take Y = 1 (through crack in a wide plate).

1. a_c = (1/π)(K_IC/σ)² = (1/π)(24/200)² = (1/π)(0.0144) = 4.58×10⁻³ m = 4.6 mm.
2. For 2024-T3 (K_IC ≈ 37 MPa√m) at the same stress: a_c = (1/π)(37/200)² = (1/π)(0.034225) = 1.089×10⁻² m = 10.9 mm.
3. The tougher 2024 tolerates a crack more than twice as long before fast fracture — and that, not its lower yield strength, is why fuselage and lower-wing skins use 2024 while 7075 goes where compression rules.
4. Inspection interval is then set so that a crack detectable by eddy current (say 1.5 mm) cannot grow to 10.9 mm within the interval, using a Paris-law crack-growth integration.

Where it is used

• Airframe primary structure: CFRP wing boxes, centre wing box, fuselage barrels and empennage — Boeing 787 (~50% composite by weight), Airbus A350 (~53%), LCA Tejas (composite wing, fin and fuselage sections built at HAL/ADA — among the highest composite fractions of any fighter), HTT-40 and Saras components at HAL/NAL Bengaluru.
• Space systems: maraging-steel and composite solid motor cases, carbon-phenolic nozzle liners, ablative heat shields, composite payload fairings and interstages, COPVs and satellite CFRP honeycomb decks — VSSC Thiruvananthapuram, LPSC, URSC Bengaluru, and vendors like Godrej Aerospace and Walchandnagar.
• Gas turbine hot section: Inconel 718 and single-crystal nickel blades, Ti fan blades and discs, CMC shrouds — HAL Engine Division Bengaluru, GTRE Kaveri, Safran/GE/Rolls-Royce supply chains in Hyderabad and Bengaluru.
• Helicopter and UAV structure: composite rotor blades with elastomeric bearings (HAL Dhruv, LCH Prachand), full-composite airframes on ADE's Rustom/TAPAS and small UAV industry.
• Job roles: stress and materials engineer, composite process/manufacturing engineer, NDT Level II/III (ASNT or NAS-410), repair and MRO engineer, quality engineer handling AS9100/NADCAP audits, and R&D at DMRL/NAL/VSSC.
• Spill-over industries: wind-turbine blades (Suzlon, LM Wind Power, Vestas plants in Tamil Nadu), Vande Bharat and metro interiors, FRP pressure vessels, automotive and sports goods — the largest employers of Indian composites graduates outside aerospace.

Interview questions

• Q: Why is specific strength more important than strength in aerospace? A: Flight performance and payload depend on structural mass, so the useful measure is strength per unit density; a strong but heavy material forces a heavier structure and less payload.
• Q: Steel is three times stiffer than aluminium — why are aircraft skins not steel? A: Because steel is also three times denser, so E/ρ is nearly identical (~26 MN·m/kg); steel buys no stiffness per kilogram, and thin steel skins would also corrode and buckle.
• Q: Where would you use 2024-T3 and where 7075-T6? A: 2024-T3 where fatigue and toughness dominate — fuselage and lower wing skins; 7075-T6 where static strength in compression dominates — upper wing skins, spars and machined fittings.
• Q: Why must fasteners in carbon composite be titanium? A: Carbon fibre is cathodic to aluminium and cadmium-plated steel; the resulting galvanic cell corrodes the fastener. Titanium is close to carbon on the galvanic series and corrosion-resistant.
• Q: What does the matrix actually do in a composite? A: It transfers load into and between fibres by shear, holds the fibre alignment, resists buckling of fibres in compression, and protects them from moisture and abrasion — and it sets the composite's maximum service temperature.
• Q: What is delamination and why is it the critical failure mode? A: Separation between plies, driven by interlaminar tension and shear where the composite has only matrix-level strength; it is caused by impact, free edges or ply drop-offs and can collapse compression strength without any visible surface damage.
• Q: What is BVID and how does it affect design? A: Barely Visible Impact Damage — internal delamination from an impact that leaves no reliably detectable surface mark; design allowables are set from compression-after-impact tests rather than pristine strength, which costs much of the composite's theoretical advantage.
• Q: Why must a laminate be balanced and symmetric? A: Symmetry makes the B matrix zero, removing bending–extension coupling and cure-induced warpage; balance (equal +θ and −θ plies) removes extension–shear coupling so an axial load does not produce shear distortion.
• Q: What does a quasi-isotropic layup like [0/±45/90] achieve? A: In-plane stiffness that is the same in every direction, at roughly one-third of the unidirectional value — used where load direction is uncertain or biaxial.
• Q: Why is the autoclave used, and what replaces it? A: Pressure (6–7 bar) plus vacuum consolidates plies, suppresses voids and raises fibre volume fraction to 0.6 for primary-structure quality; out-of-autoclave routes are RTM, VARTM/infusion and OOA prepregs, used where the part is too large or the cost too high.
• Q: Which NDT method would you choose for a suspected delamination in a 6 mm CFRP skin? A: Pulse-echo ultrasonic C-scan or phased-array UT, which maps the depth and planform of the disbond; a tap test only screens, and X-ray is better for honeycomb water ingress than for flat delamination.
• Q: Why are turbine blades made as single crystals? A: Grain boundaries transverse to the stress are the weakest path for creep and are where cavitation starts; directional solidification removes transverse boundaries and a single crystal removes them altogether, raising creep life and allowing higher turbine entry temperature.
• Q: Who certifies an Indian military aircraft structure and under what document? A: CEMILAC grants airworthiness certification with DGAQA doing quality assurance; civil aircraft and their repairs fall under DGCA's Civil Aviation Requirements.

Common mistakes

• Quoting pristine unidirectional composite strength as the design allowable — real allowables come from CMH-17-style hot/wet, compression-after-impact and B-basis statistics, typically a third or less of the UD value.
• Mixing up volume fraction and weight fraction in rule-of-mixtures problems; always convert using ρ_c = ρ_f V_f + ρ_m V_m.
• Applying the longitudinal rule of mixtures E₁ = E_f V_f + E_m V_m to the transverse direction, instead of the inverse (Reuss) form.
• Treating a laminate as isotropic and using a single E in beam or plate formulae, instead of the A, B, D matrices or at least laminate-equivalent properties.
• Forgetting the galvanic couple between carbon fibre and aluminium — no isolation ply, no sealant, and corrosion follows.
• Confusing precipitation hardening with work hardening, or quoting the wrong temper (T6 vs T73) when stress-corrosion resistance is required.
• Unit slips in specific-property and K_IC problems: MPa vs Pa, mm vs m, and forgetting that K_IC is in MPa√m so a must be in metres.
• Saying composites "do not fail in fatigue" — they are excellent in the fibre direction but matrix-dominated, off-axis and bonded-joint fatigue is real, and moisture plus temperature accelerate it.

Related topics

Aircraft structures · Finite element analysis basics · Mechanics of materials and failure theories · Fracture mechanics and fatigue · Gas turbine and propulsion materials · Manufacturing processes and NDT.