Airframe Study Guide

Sheet Metal Repair

Sheet metal repair is heavily tested on the FAA Airframe written exam. Master rivet selection, damage classification, and repair procedures to score well in this category.

1. Types of Aircraft Sheet Metal

The most common aluminum alloy used for aircraft skin is 2024-T3. It offers high strength, good fatigue resistance, and excellent machinability. The “T3” designator means the alloy was solution heat treated and then cold worked (strain hardened). It is used for fuselage skin, wing skin, and primary structural panels.

6061-T6 is a more corrosion-resistant alloy used for less-structurally-critical parts such as ribs, brackets, and fairings. It is easier to weld than 2024 but has lower fatigue strength, so it is not used for primary skin panels on transport category aircraft.

Key property comparison:

2024-T3: High strength, moderate corrosion resistance — primary structural skin
6061-T6: Good corrosion resistance, weldable — secondary structure and fittings
7075-T6: Highest strength of common alloys — spars, highly loaded fittings

2. Damage Classifications

When an aircraft sustains damage to sheet metal, the mechanic must classify the damage to determine the correct course of action. Per AC 43.13-1B and the aircraft's Structural Repair Manual (SRM), damage falls into three categories:

Negligible Damage — Small scratches, dents, or nicks that can be polished out or left as-is without structural concern. No repair needed. Damage limits for negligible damage are defined in the SRM.
Repairable Damage — Damage that exceeds negligible limits but can be repaired by stop-drilling (for cracks), patch repair, or doubler installation. The SRM specifies the maximum repairable damage size.
Beyond Limits (Replacement Required) — Damage that exceeds the maximum repairable damage limit. The affected skin panel, stringer, or frame must be replaced. Only approved replacement parts from the manufacturer may be used.

3. Stop-Drilling Cracks

Cracks in sheet metal propagate because of stress concentration at the crack tip. The standard technique to arrest crack growth is stop-drilling: drilling a small hole at the tip of the crack to eliminate the sharp stress concentration and distribute the stress over a larger area.

The standard stop-drill size is #30 drill (approximately 1/8 inch). The hole must be centered precisely at the crack tip — not beyond it. After stop-drilling, the area must be evaluated to determine whether a patch or doubler repair is also required.

Stop-drill procedure:

Dye penetrant or magnification to find exact crack tip location
Mark the center of the stop-drill hole at the crack tip
Drill with a #30 (0.128″) drill bit, deburr both sides
Inspect for additional cracks before closing the area

4. Rivet Selection Rules

Choosing the correct rivet is critical. The wrong diameter or inadequate edge/pitch distance will result in a weak or failed repair. The key rules from AC 43.13-1B:

Diameter rule: Rivet diameter = 3 × thickness of the thickest sheet. Round up to the nearest standard diameter (3/32, 1/8, 5/32, 3/16 inch).
Minimum edge distance: 2× rivet diameter (2D) from the center of the rivet to the edge of the material. Recommended is 2.5D.
Minimum rivet pitch (spacing): 3× rivet diameter (3D) center-to-center. Recommended pitch is 4D to 6D.
Head type: AN426 (100° flush/countersunk) for aerodynamic surfaces. AN470 (universal/round head) for non-aerodynamic or structural repairs.

5. Corrosion Treatment

Corrosion is one of the most common causes of structural damage in aircraft. The treatment procedure follows a defined sequence:

Identify the type of corrosion: surface (white/gray powder on aluminum), intergranular (along grain boundaries), galvanic (dissimilar metals in contact), or pitting.
Remove all corrosion products. For aluminum: use aluminum wool, Scotch-Brite, or a chemical conversion coating process. Never use steel tools on aluminum — they will cause galvanic contamination.
Treat the cleaned surface with a chemical conversion coating (Alodine/Iridite) to restore corrosion resistance without removing base material.
Seal and protect with an epoxy primer and topcoat. Seal all seams with approved sealant to prevent moisture intrusion.

6. Bend Allowance

When sheet metal is bent, the outer surface is stretched and the inner surface is compressed. The neutral axis — located approximately 0.445× the material thickness from the inside — is the line that neither stretches nor compresses. Bend allowance is the arc length along the neutral axis.

The bend allowance formula: BA = (0.01743 × R + 0.0078 × T) × degrees where R is the inside bend radius and T is the material thickness. For exam purposes, bend allowance tables are usually provided.

Use the Bend Allowance Calculator to calculate setback and flat pattern dimensions for your repair.

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