| 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).
