composite-aerostructures

name	composite-aerostructures
description	Composite aerostructures — laminate design for wings/fuselage, certification (FAR 25 / CS-25), damage tolerance (CAI, BVID), environmental knockdowns, repair, sandwich panels, autoclave process.
metadata	{"priority":7,"promptSignals":{"phrases":["composite aerostructure","composite aircraft","BVID","compression after impact","CAI","FAR 25 composite","composite wing"],"minScore":3}}

Composite Aerostructures — Complete Skill

FAR/CS-25 Composite Certification Framework (AC 20-107B)

Design Levels

Limit load: maximum load expected in service; no permanent deformation Ultimate load: limit × 1.5; no failure; positive margin of safety required

Static strength: ultimate load must be carried at ambient and environmental conditions (hot-wet)

Damage tolerance: Structure must sustain limit load with damage (Barely Visible Impact Damage — BVID) Structure must sustain 70% limit load (some interpretations: 83%) with visible damage

BVID Definition

BVID = barely visible impact damage (not reliably detected by walk-around inspection) Impact energy threshold: 6 ft·lbf (8 J) standard (FAA policy); 30 J for critical areas Maximum damage size at BVID threshold: typically 7.5–25 mm diameter

Requirement: residual strength ≥ ultimate load after BVID exposure

Compression After Impact (CAI)

Impacts create delamination → dramatic reduction in compression strength CAI test: ASTM D7137 (formerly D6264); 152×102 mm specimen; drop weight impact Typical CAI strength: 60–80% of undamaged compression strength (after BVID level impact)

CAI strength formula (empirical): F_cai ≈ F_c0 × (1 - α √(d_damage / t)) α ≈ 0.3–0.5; d_damage = damage diameter; t = laminate thickness

Design approach: F_allow = F_cai / (KT × environment factor × scatter factor) KT = load redistribution factor; environment factor = B-basis hot-wet value

Environmental Knockdowns

Moisture and temperature degrade matrix-dominated properties: Hot-wet condition: Tg - 25°C (design temperature = wet Tg minus 25°C margin) Tg (dry) of standard epoxy: 177°C (350°F cure); 121°C (250°F cure) Tg (wet, at 85% RH) ≈ Tg (dry) - 20–30°C

Environmental knockdown factors (B-basis, typical): Tension (fiber-dominated): 0.95–1.00 (small effect) Compression (matrix-dominated): 0.75–0.85 (significant) Shear: 0.70–0.80 Interlaminar shear: 0.65–0.75

Design allowables: A-basis: 99% reliability, 95% confidence (single load path) B-basis: 90% reliability, 95% confidence (redundant load path, most airframe)

Laminate Design Guidelines (Aerospace)

Stacking sequence rules (per Boeing/Airbus DTS):

• 10% rule: no less than 10% plies in each of 0°, ±45°, 90° orientations
• Balanced: equal +θ and -θ plies (prevents shear coupling)
• Symmetric: stacking symmetric about mid-plane (prevents bending-extension coupling)
• Avoid grouping more than 4 plies of same orientation (delamination risk)
• 0° plies at outer surfaces for edge damage protection
• 45° plies adjacent to 0° plies for interlaminar shear

Typical laminate fractions: Primary tension/compression structure: 60% 0°, 30% ±45°, 10% 90° Shear panels (wing skin): 40% 0°, 50% ±45°, 10% 90° General structure: 25% each of 0°, ±45°, 90° (quasi-isotropic)

Sandwich Structure

Core materials: Nomex honeycomb: 48 kg/m³ density; G_c ≈ 50–100 MPa; compression ≈ 2 MPa Aluminum honeycomb: denser; higher shear; used in high-load control surfaces PMI foam (Rohacell): uniform; closed cell; no moisture ingress; 52–110 kg/m³

Sandwich design: Flexural stiffness: D ≈ E_f t_f d² / 2 (face sheet contribution; t_f = face, d = separation) Core shear: τ_c = Q / (b d) (must be < G_c/SF) Face sheet wrinkling: σ_wr = 0.5 (E_f E_c G_c)^(1/3)

Manufacturing — Autoclave

Prepreg layup: 125°C / 350°F systems (most aerospace) Cure cycle: ramp to 177°C; hold 90–120 min; 6–7 bar (90 psi) pressure Fiber volume fraction V_f: target 55–62%; controlled by bleeder and pressure Void content: ≤ 1% required for primary structure; ≤ 2% for secondary

Out-of-autoclave (OoA): Prepreg designed for vacuum bag only cure; V_f and voids slightly higher Reduces cost; suitable for large structures (B787 fuselage sections: 28 m long)

Repair Classification (AC 43.13-1B / SRM)

Cosmetic: no structural function impaired; filler and paint Allowable damage: below threshold; no repair needed; document Repairable: within Structural Repair Manual limits; patch repair Allowable damage limit exceeded: depot repair or part replacement

Wet layup repair: room-temperature or elevated cure; for non-critical structure Prepreg patch repair: higher performance; hot bonded; for primary structure

Inspection Methods

NDT for composites: Ultrasonic (pulse-echo, TTU): delaminations, disbonds Thermography (flash IR): subsurface damage; foreign objects; BVID indication Tap testing: simple; manual; disbonds in thin laminates X-ray (digital): foreign objects; 2D fiber waviness CT scan: most comprehensive; 3D void and damage mapping

Output

Provide: laminate stacking sequence, design allowables (B-basis, hot-wet) [MPa], CAI strength [MPa], BVID impact energy threshold [J], environmental knockdown applied, margin of safety at ultimate and limit load, sandwich core/face parameters, repair category assessment.