airframe-fatigue

name	airframe-fatigue
description	Airframe fatigue — safe-life/damage-tolerant/fail-safe design philosophies, spectrum loading, crack growth da/dN, AFGROW/NASGRO, gust spectrum (FAR 25.571), structural test articles, MIL-STD-1530.
metadata	{"priority":7,"promptSignals":{"phrases":["airframe fatigue","aircraft fatigue","damage tolerance aircraft","safe life design","fail safe structure","fatigue spectrum aircraft","FAR 25.571"],"minScore":3}}

Airframe Fatigue — Complete Skill

Design Philosophies

Safe Life

Concept: structure designed to survive without crack initiation for entire service life; retired at service life Life limit: L_safe = L_test / scatter factor (SF = 4–6 for aircraft) When used: landing gear, fittings, helicopter dynamic components (MIL-STD-1530) Testing: full-scale fatigue test to 4× design service objective

Damage Tolerant

Concept: accept that cracks will exist; ensure crack will not grow to critical size within inspection interval Key equation: a_crit = 1/π × (K_IC / (σ × F))² [m; critical crack size; K_IC = fracture toughness; F = geometry factor; σ = operating stress] Inspection interval: t_inspect ≤ (N_crit - N_init) / 2 [must be detected at least 2 inspection intervals before fracture]

Residual strength: P_residual ≥ 1.0 P_limit [can carry limit load with crack at critical size]

Fail Safe

Concept: multiple load paths; crack/failure of one element does not cause catastrophic failure Design: redundant structure; arrested crack propagation; visible damage Used for: commercial fuselage panels, wing skins (multiple stringers)

Loading Spectra

Gust spectrum (FAR 25.571, Appendix A): Gust alleviation factor: K_g = 0.88μ_g / (5.3 + μ_g) μ_g = 2m / (ρ c S C_Lα) [mass ratio; m = aircraft mass; c = mean chord; S = wing area; ρ = air density] Derived gust velocity: U_de = U_ref × F_g [F_g = flight profile factor]

Standard gust spectra (Pratt spectrum): Δn per gust = K_g × U_de × a × V / (2 W/S) [n = load factor increment; a = lift curve slope; W/S = wing loading; V = airspeed]

Taxi spectrum: sinusoidal bump spectrum from runway; critical for landing gear and fuselage lower structure Maneuver spectrum: g-exceedance curves per AFDD; 3.5 g → 1/10⁴ flights typical for fighter

Rainflow counting: extract load cycles from time history n_i = number of cycles at stress range Δσ_i from rainflow counting

Crack Growth (Fracture Mechanics)

Paris law: da/dN = C × (ΔK)^m [m/cycle; ΔK = stress intensity factor range [MPa√m]] ΔK = Δσ × F(a, geometry) × √(πa) [F = geometry factor from handbooks]

Paris law constants (2024-T3 aluminum): C = 4.0×10⁻¹¹; m = 3.3 [metric; da/dN in m/cycle; ΔK in MPa√m]

NASGRO equation (more complete): da/dN = [C × (1 - f)^n × ΔK^n × (1 - ΔK_th/ΔK)^p] / [(1 - R)^n × (1 - K_max/K_IC)^q] f = crack opening ratio (Newman); ΔK_th = threshold; R = stress ratio; C, n, p, q = material constants

Stress intensity solutions: Through crack in infinite plate: K = σ √(πa) Edge crack: K = 1.12 σ √(πa) Corner crack at fastener hole: K = 2.0 σ √(πa) (approximate)

Software: AFGROW and NASGRO

AFGROW (USAF Research Laboratory, free): Library of geometry solutions; retardation models (Wheeler/Willenborg); spectrum loading Input: geometry (crack shape, location), material (Paris constants), load spectrum Output: a vs. N curves; critical crack size; inspection intervals

NASGRO (NASA/FAA/ESA, commercial): Includes Willenborg retardation; large geometry library; crack closure Certified for FAA/EASA regulatory use

Critical Structure Fatigue Analysis Process (MIL-STD-1530D)

Step 1 — Stress analysis: FEM or hand calculation; identify principal stresses at critical locations (fastener holes, cutouts, joints) Stress concentration factor Kt; gross section → net section stress

Step 2 — Load spectrum development: Spectrum from usage data or design usage (MIL-HDBK-516C) Rainflow count applied to load-time history → exceedance histogram

Step 3 — Material selection and allowables: A-basis or B-basis allowables (MIL-HDBK-5/MMPDS) For damage tolerance: K_IC, ΔK_th, Paris constants

Step 4 — Crack growth analysis: From assumed initial crack (NDE detection threshold) to critical Select inspection threshold: a_0 = 1.27 mm (0.05 in) for standard inspections

Step 5 — Inspection interval: t_inspect = (N(a_detect) - N(a_0)) / 2 a_detect = reliably detected crack size for inspection method

Inspection Methods for Airframe

Method	Min detectable crack	Application
Visual	10–25 mm	Skin surfaces
Eddy current	0.5–1.5 mm	Surface/sub-surface
Ultrasonic (UT)	1–3 mm	Thick structure
X-ray	1–5 mm	Complex structure
Dye penetrant	0.1–0.5 mm	Surface access required

Enhanced Zonal Inspection Program (EZAP): systematic approach per MSG-3 (ATA); defines inspection zones and thresholds

Full-Scale Fatigue Testing

Duration: 2× to 4× design service objective (DSO) Load application: hydraulic jacks simulate flight + ground loads; wing tip deflection; cabin pressurization cycles Data: strain gauges at 500–2000 locations; visual + NDI inspection every 1000 cycles

Fail-safe testing: cut one element (stringer, frame) at end of test; verify residual strength ≥ limit load

Standards

Standard	Scope
FAR/CS 25.571	Damage-tolerance/fatigue evaluation
MIL-STD-1530D	Aircraft structural integrity program
MIL-HDBK-516C	Airworthiness criteria
MMPDS (was MIL-HDBK-5)	Metallic material allowables
AC 91-56B	Structures continued airworthiness

Output

Provide: design philosophy (safe-life/damage-tolerant/fail-safe), critical location stress [MPa] and Kt, Paris law constants (C, m), initial crack assumption a_0 [mm], critical crack size a_crit [mm], crack growth life from a_0 to a_crit [cycles or hours], inspection interval [hours], inspection method and minimum detectable crack size [mm], residual strength with critical crack [MPa], full-scale test requirement (DSO × n), scatter factor applied, and applicable standard (FAR 25.571, MIL-STD-1530D).