5.1 Metallic Structure Inspection and Damage Patterns

How Airframe Structure Carries Load

Almost every metal airplane uses stressed-skin (monocoque or semimonocoque) construction. In a pure monocoque structure the external skin alone carries the loads — there is no internal bracing — so any skin dent or crack is structurally significant and the design is heavy for its strength.

Because monocoque skins must be thick to resist buckling, aircraft instead use semimonocoque construction almost universally: the skin still carries load, but it is supported by longerons and stringers (running lengthwise) and frames, formers, and bulkheads (running around the cross-section). This lets the skin be thin while the substructure handles bending and keeps the shape.

The major airframe assemblies are the fuselage (longerons, stringers, frames, bulkheads, skin), the wing (front and rear spars, ribs, stringers, and skin), and the empennage (the vertical and horizontal stabilizers with their own spars and ribs). A bulkhead is a heavy frame that divides the fuselage or anchors major loads such as a wing attach or pressurization dome. Knowing what part you are looking at tells you the load it carries and therefore how seriously to treat damage.

The Five Stresses and What Damage Reveals

Structure must resist five basic stresses, and damage patterns point to which one is being overworked:

Most rivets and bolts in a sheet-metal joint are loaded in shear, which is why fastener selection and edge distance matter. A wing spar in bending has its cap in tension on one face and compression on the other, so a crack on the tension side grows fast. Reading a damage scenario as a load-path question — not just a cosmetic one — is the core skill the Airframe test rewards.

Inspecting and Classifying Defects

Nondestructive inspection (NDI) method must match the material and the suspected defect:

• Visual / magnification / borescope — first line for cracks, corrosion, loose fasteners.
• Dye (liquid) penetrant — surface-breaking cracks in nonporous, nonmagnetic metals (aluminum, magnesium, some stainless).
• Eddy current — surface and near-surface cracks in conductive metals; great around fastener holes.
• Magnetic particle — surface/subsurface flaws in ferromagnetic (steel) parts only — it will NOT work on aluminum.
• Ultrasonic — subsurface flaws and thickness.
• Radiographic (X-ray) — hidden internal defects and corrosion, with strict safety controls.

Classify what you find: a crack grows under cyclic load and is normally serious; a smoking rivet (dark teardrop streak) or fretting indicates a working/loose joint; an elongated hole has lost bearing area and clamp-up; a scratch is judged by depth, direction, and location; heat damage can alter aluminum temper even when the shape is normal. A repair material is acceptable only when alloy, temper, thickness, grain direction, and fastener compatibility match the approved data — strength alone is not enough.

A Disciplined Decision Sequence

Use the same logic on every metallic-structure scenario:

1. Identify the structure and its likely load path (tension/compression/shear/bending/torsion).
2. Clean enough to inspect without needlessly stripping protective finish.
3. Classify the defect type and measure it.
4. Compare to the Structural Repair Manual (SRM), aircraft maintenance manual (AMM), service bulletins, and airworthiness directives.
5. Decide whether the damage is within published negligible/repairable limits or requires approved engineering data (a DER/8110-3 or manufacturer approval).
6. Document the finding, the data used, and the result.

Sheet-metal work also creates hazards the ACS risk elements expect you to manage: burrs cut hands and start cracks, drill chips fly into eyes (wear eye protection), compressed air and pneumatic tools must be controlled, and an incompatible repair material can set up galvanic corrosion. Exam questions frequently hide the key in the location of the damage — a mark near a spar cap, bend radius, bulkhead, or hinge deserves far more caution than the same mark on an unstressed fairing.

When in doubt, choose the answer that inspects first, uses current approved data, preserves temper and load path, and never returns a part to service on appearance alone.

Common Exam Traps and Worked Logic

Several recurring traps appear in metallic-structure questions. First, strength is not the only criterion for a repair material: a replacement skin must match the original alloy and temper (for example, 2024-T3 versus 7075-T6), the correct thickness or one gauge thicker if the data allows, the grain direction when forming bend radii, and fastener compatibility to avoid galvanic action. A part that is merely 'as strong' but a different alloy can corrode, fatigue, or fail to flex with the surrounding structure.

Second, negligible damage is defined by the data, not by feel — the SRM states the exact dent depth, scratch length, and edge-distance reductions that may be left unrepaired or simply blended out and re-protected.

A short worked case shows the load-path mindset. 040-inch lower wing skin (a tension-loaded surface in flight) has a crack running spanwise from a fastener hole. Because the lower skin is in tension during normal flight, the crack will propagate under cyclic load; this is not a candidate for a quick cosmetic fix.

The correct sequence is to stop-drill the crack tip only if the SRM authorizes it as an interim step, inspect the surrounding holes with eddy current for additional cracking, and then install the manufacturer's approved patch — typically a doubler that re-establishes the rivets the crack interrupted, in the same alloy and gauge. Choosing eddy current (not magnetic particle) is itself a tested point, because the skin is aluminum.

The disciplined order — identify load path, classify, inspect with the right method, repair only within approved limits, document — is exactly what distinguishes the airworthy answer from the tempting shortcut.