1. ROLE

3. GETTING STARTED

When you trigger this skill, I'll work with whatever you give me — but the more context, the better the output.

Minimum I need (pick one):

• "Check if a 2 mm Al 7075 panel can survive Falcon 9 launch loads"
• "Select materials for a 300 kg satellite primary structure"
• "What's the buckling load of a 1.2 m diameter cylindrical adapter?"

Helpful if you have it:

• Spacecraft total mass and center of gravity location
• Launch vehicle (defines coupled loads environment)
• Structural configuration (monocoque, skin-stringer, sandwich, truss)
• Design life and mission environment (LEO, GEO, deep space)
• Temperature range (affects allowables)
• Heritage constraints (flight-proven materials, processes)

What I'll ask if you don't specify:

• "What launch vehicle?" — load environment changes everything
• "Primary or secondary structure?" — determines safety factors and load paths
• "Mass target or mass budget?" — drives material selection aggressively
• "Temperature range?" — cryogenic, ambient, or high-temp changes allowables

4. CONNECTORS

Shared Tools (in shared/tools/)

Shared Data (in shared/ — pack-level)

Cross-skill Connectors

5. TAXONOMY

5.1 Aerospace Structural Materials

*CFRP properties depend on layup; values shown for quasi-isotropic to unidirectional range.

5.2 Specific Strength Comparison

5.3 Launch Load Environments

5.4 Launch Vehicle Loads Summary

6. PROCESS

Step 1: Define Load Environment

• Launch vehicle → user manual defines quasi-static, sine, random, acoustic, shock
• Combined load case → axial and lateral act simultaneously
• Design load = limit load × uncertainty factor (typically 1.25 per ECSS)
• IF launch vehicle not specified → ASK. Load environment changes everything.
• IF multiple launchers → use envelope of worst case from each.

Step 2: Free-Body Diagram & Load Path

• Identify load path: launch vehicle adapter → primary structure → secondary structure → equipment
• Interface loads: F_axial = m × n_axial × g₀, F_lateral = m × n_lateral × g₀
• Combined: F_total = sqrt(F_axial² + F_lateral²) for preliminary sizing
• Moments at interfaces: M = F_lateral × L_cg (distance from interface to CoG)

Step 3: Stress Analysis

Direct stress:

σ = F / A                         (axial stress)
σ_b = M·y / I                    (bending stress)
τ = V·Q / (I·t)                  (shear stress)
σ_vm = sqrt(σ² + 3τ²)           (von Mises equivalent stress)

Thin-wall cylinder under axial + bending:

σ_max = F/(2πRt) + M/(πR²t)     (compression side governs)

Step 4: Buckling Analysis

Euler column buckling:

P_cr = π²EI / L²                 (simply supported)
σ_cr = π²E / (L/r)²              (in terms of slenderness ratio)

Thin-wall cylinder buckling (with NASA SP-8007 knockdown):

σ_cr_classical = 0.605 × E × t/R
σ_cr_design = γ × σ_cr_classical      (γ = knockdown factor, typically 0.3–0.65)

Knockdown factor γ depends on R/t ratio:

Step 5: Margin of Safety Calculation

MS_yield    = σ_allow_yield / (FS_yield × σ_applied) - 1
MS_ultimate = σ_allow_ultimate / (FS_ultimate × σ_applied) - 1

ECSS-E-ST-32C Safety Factors:

Rule: MS ≥ 0 required. MS < 0 = structural failure risk. MS > 0.5 = probably over-designed (mass penalty).

Step 6: Worked Example — Satellite Primary Structure Panel

Problem: A 400 kg satellite launches on Falcon 9. Check the central cylinder (Al 7075-T6, R = 0.5 m, t = 2.0 mm, L = 0.8 m) under combined axial + bending launch loads.

Given:

• Satellite mass: m = 400 kg
• Falcon 9 loads: n_axial = 6.0 g, n_lateral = 2.0 g
• Material: Al 7075-T6 (σ_y = 503 MPa, σ_u = 572 MPa, E = 71.7 GPa)
• Cylinder: R = 0.5 m, t = 2.0 mm, L = 0.8 m
• CoG height above interface: L_cg = 0.6 m

Step A — Loads:

F_axial   = 400 × 6.0 × 9.81 = 23,544 N
F_lateral = 400 × 2.0 × 9.81 =  7,848 N
M_base    = F_lateral × L_cg  =  7,848 × 0.6 = 4,709 N·m

Step B — Stress in cylinder wall (compression side):

A = 2π × R × t = 2π × 0.5 × 0.002 = 6.283 × 10⁻³ m²
I = π × R³ × t = π × 0.125 × 0.002 = 7.854 × 10⁻⁴ m⁴
σ_axial  = F / A           = 23,544 / 6.283×10⁻³  =  3.75 MPa
σ_bending = M×R / I        = 4,709 × 0.5 / 7.854×10⁻⁴ =  3.00 MPa
σ_combined = 3.75 + 3.00   =  6.75 MPa (compression side)

Step C — Buckling check:

σ_cr_classical = 0.605 × E × t/R = 0.605 × 71,700 × 0.002/0.5 = 173.5 MPa
R/t = 0.5 / 0.002 = 250  →  γ = 0.50
σ_cr_design = 0.50 × 173.5 = 86.8 MPa

Step D — Margins of Safety:

MS_yield    = 503 / (1.25 × 6.75) - 1 = 503 / 8.44 - 1 = 58.6  ✓ (very high)
MS_ultimate = 572 / (1.50 × 6.75) - 1 = 572 / 10.13 - 1 = 55.5  ✓ (very high)
MS_buckling = 86.8 / (1.50 × 6.75) - 1 = 86.8 / 10.13 - 1 = 7.57  ✓ (positive)

Step E — Verdict: All margins positive. Buckling governs (MS = 7.57) over strength (MS > 55). Structure is significantly over-designed for this mass — could reduce thickness to 1.0 mm (MS_buckling ~ 3.3) or switch to Al 6061-T6 to save cost. Structural mass of cylinder: ρ × A × L = 2810 × 6.283×10⁻³ × 0.8 = 14.1 kg.

7. OUTPUT TEMPLATE

# [Mission Name] — Structural Analysis Report

## Configuration
| Parameter | Value |
|-----------|-------|
| Spacecraft mass | [X] kg |
| Launch vehicle | [name] |
| Structure type | [monocoque/sandwich/truss] |
| Primary material | [alloy/composite] |

## Load Environment
| Load Case | Axial (g) | Lateral (g) | Combined (N) |
|-----------|-----------|-------------|-------------|
| Quasi-static | [X] | [X] | [X] |
| Design (×1.25) | [X] | [X] | [X] |

## Stress Summary
| Element | σ_applied (MPa) | σ_allow (MPa) | FS | MS | Status |
|---------|-----------------|---------------|-----|-----|--------|
| [component] | [X] | [X] | [X] | [X] | [PASS/FAIL] |

## Buckling Summary
| Element | σ_cr (MPa) | σ_applied (MPa) | γ | MS | Status |
|---------|-----------|-----------------|---|-----|--------|
| [component] | [X] | [X] | [X] | [X] | [PASS/FAIL] |

## Mass Budget
| Component | Mass (kg) | % of total |
|-----------|-----------|-----------|
| Primary structure | [X] | [X]% |
| Secondary structure | [X] | [X]% |
| Mechanisms | [X] | [X]% |
| **TOTAL structural** | **[X]** | **[X]%** |

## Recommendation
[Material selection rationale, mass optimization opportunities, test recommendations]

8. CLASSIFICATION

9. VARIATIONS

• A: SmallSat (<100 kg) — Al 6061-T6 or Al 7075-T6 monocoque/tray, standard ESPA or CubeSat deployer, random vibration governs, qualification by test, structural mass 20–30% of total
• B: Large GEO Satellite (3000–6000 kg) — CFRP sandwich central tube, Al honeycomb panels, CLA with launch vehicle, sine vibration governs, notching at resonances, structural mass 10–15% of total
• C: Deployable Structures — Solar arrays, antenna reflectors, booms; deployment shock + thermal snap loads; latch preload > 3× release load; stiffness: f₁ > 0.1 Hz deployed, f₁ > 40 Hz stowed
• D: Pressurized Structures — Propellant tanks, pressure vessels; MEOP × 1.5 yield FS, × 2.0 ultimate FS; leak-before-burst criterion; proof test to 1.5× MEOP mandatory
• E: Re-entry Vehicles — Aerothermal + mechanical loads combined; TPS bond analysis; ablation mass loss; thermal stress from 200°C+ gradient; fracture mechanics for safe-life

10. ERRORS & PITFALLS

• E1: Ignoring combined loads — axial and lateral act simultaneously, not sequentially; use root-sum-square or coupled loads analysis, never just the worst single axis
• E2: Using room-temperature allowables at extreme temperatures — Al loses 50% strength above 200°C; CFRP matrix degrades above 120°C; cryogenic embrittlement of some steels
• E3: Forgetting knockdown factors for buckling — classical Euler/shell buckling overpredicts by 2–5× for real cylinders; always apply NASA SP-8007 or ECSS knockdown
• E4: Stiffness vs. strength confusion — spacecraft often fail stiffness requirements (f₁ > 100 Hz stowed) long before stress limits are reached; check frequency first
• E5: CTE mismatch at interfaces — CFRP (-0.5 µm/m·K) bonded to Al (23.6 µm/m·K) with ΔT = 150°C generates 3.6‰ strain → delamination; always check thermal stress at joints
• E6: Ignoring fitting factors — ECSS requires 1.15 additional FS on all mechanically fastened joints; forgetting this turns MS = +0.10 into MS = -0.05
• E7: Optimistic mass estimates — "structure is 8% of total" works for large satellites; SmallSats are 20–30%; CubeSats can be 35%; always use class-appropriate parametrics
• E8: Not accounting for insert pull-out — threaded inserts in honeycomb sandwich have allowable pull-out of 500–2000 N depending on core density; this often governs equipment mounting

11. TIPS

• T1: Start from stiffness, not strength — spacecraft frequency requirement (f₁ > 100 Hz) usually sizes structure before stress does
• T2: Buckling governs thin-walled design — for R/t > 200, buckling stress is 10–50× lower than yield; always check buckling first
• T3: Use sandwich panels for large flat areas — 10× stiffness of solid panel at 20% of the mass; honeycomb core density 32–96 kg/m³
• T4: Structural mass fractions by class — SmallSat 20–30%, medium 12–18%, large GEO 10–15%, interplanetary 15–20%
• T5: Material selection shortcut — Al 7075-T6 for machined fittings, Al 6061-T6 for welded/sheet, Ti-6Al-4V for thermal standoffs and high-load bolts, CFRP for mass-critical panels
• T6: Fastener rule of thumb — bolt circle diameter ≥ 0.6 × cylinder diameter; minimum 8 bolts for interface; preload = 0.65 × proof load for flight
• T7: Calibrate against heritage — Eurostar 3000: 4700 kg, structure 540 kg (11.5%); 1U CubeSat: 1.33 kg, structure 0.4 kg (30%); ISS truss segment: 10,000 kg, structure 4,500 kg (45%)
• T8: Test philosophy — protoflight: test to 1.25× limit (qual level, acceptance duration); qualification: test to 1.5× limit with dedicated unit; random vib: 2 min/axis protoflight, 3 min/axis qual

12. RELATED SKILLS