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):

• "Size a solar array for a 500 W payload in LEO"
• "How big a battery for 35 minutes of eclipse at 200 W?"
• "Power budget for a 6U CubeSat with an X-band transmitter"

Helpful if you have it:

• Orbit altitude and inclination (drives eclipse duration and radiation)
• Payload power demand by mode (nominal, peak, standby)
• Mission lifetime (years — drives degradation)
• Mass or area constraints (stowage volume in fairing)
• Temperature range at array (hot/cold case)
• Pointing strategy (sun-tracking vs body-mounted)

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

• "What orbit? LEO/SSO/GEO/deep space?" — eclipse fraction changes 10x
• "Mission lifetime?" — 1 year vs 15 years changes array size by 30%+
• "Sun-tracking or body-mounted arrays?" — factor of 3 in required area

4. CONNECTORS

Shared Tools (in shared/tools/)

Shared Data (in shared/ — pack-level)

Cross-skill Connectors

5. TAXONOMY

5.1 Solar Cell Types

Standard reference condition: AM0, 1361 W/m², 28 °C cell temperature.

5.2 Battery Types

LEO = ~5,400 cycles/year. GEO = ~90 cycles/year (eclipse seasons only).

5.3 RTG and Nuclear Power Sources

RTG degradation: ~0.8%/year (Pu-238 decay) + thermocouple degradation (~1.0%/year combined).

5.4 Power Bus Architectures

Rule of thumb: Harness mass ∝ 1/V² for same power. Going from 28 V to 100 V cuts harness mass by ~13x.

6. PROCESS

Step 1: Mission Definition

• Orbit: altitude, inclination, LTAN (drives beta angle range)
• Payload power: watts by mode (standby, nominal, peak)
• Mission lifetime: years (drives EOL degradation)
• Constraints: mass limit, stowed volume, pointing strategy

IF orbit is not specified → ASK. IF payload power is not specified → provide parametric analysis for 100 W, 500 W, 1 kW, 5 kW.

Step 2: Eclipse Duration Calculation

Orbit period:          T = 2π × √(a³ / μ)
Eclipse half-angle:    ρ = arcsin(R_Earth / a)
Eclipse fraction:      f_e = ρ / π  (worst case, β = 0°)
Eclipse duration:      T_eclipse = f_e × T
Sunlit duration:       T_sun = T - T_eclipse

WORKED EXAMPLE — 525 km SSO (97.4° inclination):

a = 6371 + 525 = 6896 km
T = 2π × √(6896³ / 398600) = 5706 s = 95.1 min
ρ = arcsin(6371 / 6896) = 67.5°
f_e = 67.5 / 180 = 0.375  (worst case β = 0°)
T_eclipse = 0.375 × 95.1 = 35.7 min
T_sun = 95.1 − 35.7 = 59.4 min

Note: At β > ~71° for 525 km SSO, eclipse duration → 0 (full sun orbit). Typical SSO β range: 0° to ~23.5° + orbit plane drift.

Step 3: Power Budget by Mode

Build a power budget table for every operational mode.

WORKED EXAMPLE — 525 km SSO Earth observation satellite (800 W payload):

Orbit-average power requirement:

Assume duty cycle per orbit: 15% imaging, 15% downlink, 70% nominal
P_orbit_avg = 0.15 × 1170 + 0.15 × 313 + 0.70 × 156 = 175.5 + 47.0 + 109.2 = 331.7 W
Add 20% system margin: P_req = 331.7 × 1.20 = 398 W

Step 4: Solar Array Sizing

Master formula:

P_sa = P_req × T_orbit / (T_sun × η_path × cos(θ) × (1 − D)^L × F_packing × F_temp)

Where:

• P_req = orbit-average power requirement (W)
• T_orbit / T_sun = orbit period / sunlit time = duty correction
• η_path = power path efficiency (PCDU + harness + battery round-trip) = 0.80-0.90
• cos(θ) = worst-case sun incidence angle (sun-tracking → cos 0° = 1.0; body-mounted → cos 23.5° = 0.917)
• D = annual degradation rate, L = mission life (years)
• F_packing = cell packing factor on substrate (typically 0.85-0.90)
• F_temp = temperature derating factor (cells lose ~0.5%/°C above 28°C; at 65°C in LEO → F_temp ≈ 0.82)

WORKED EXAMPLE — 525 km SSO, 5-year mission, sun-tracking arrays, triple-junction ZTJ cells:

P_req = 398 W (from Step 3)
T_orbit / T_sun = 95.1 / 59.4 = 1.601
η_path = 0.85 (regulated bus, includes battery charge/discharge)
cos(θ) = 0.97 (sun-tracking with ±15° seasonal variation)
Cell η_BOL = 29.5%
D = 1.5%/yr → (1 − 0.015)^5 = 0.927 → EOL factor
F_packing = 0.87
F_temp = 0.83 (average operating temp ~60°C in LEO)

P_sa (at cell level) = 398 × 1.601 / (0.85 × 0.97 × 0.927 × 0.87 × 0.83)
P_sa = 637.2 / (0.85 × 0.97 × 0.927 × 0.87 × 0.83)
P_sa = 637.2 / 0.5528
P_sa = 1153 W (BOL array power needed)

Array area = P_sa / (1361 × η_cell) = 1153 / (1361 × 0.295) = 1153 / 401.5 = 2.87 m²
Array mass = 2.87 m² × 3.2 kg/m² (rigid panel) = 9.2 kg (panels only)
  + deployment mechanism ~3 kg + wiring ~1.5 kg = ~13.7 kg total

Step 5: Battery Sizing

Master formula:

C_battery = P_eclipse × T_eclipse / (DoD × V_bus × η_discharge)

WORKED EXAMPLE — same 525 km SSO satellite:

P_eclipse = 200 W (from power budget, eclipse mode with margin)
T_eclipse = 35.7 min = 0.595 hr
DoD = 25% (Li-ion, LEO, 5-year life → ~27,000 cycles needed → 25% DoD gives >50,000 cycles)
V_bus = 28 V
η_discharge = 0.95

C_battery = 200 × 0.595 / (0.25 × 28 × 0.95)
C_battery = 119 / 6.65
C_battery = 17.9 Ah

Select: 8S2P configuration of 18650 Li-ion cells (3.6 V × 8 = 28.8 V, 3.2 Ah × 2 = 6.4 Ah per string)
   → Need 3 parallel strings: 8S3P = 24 cells, 9.6 Ah × 28.8 V = 276.5 Wh
   → Actual DoD = 119 Wh / 276.5 Wh = 43% — TOO HIGH for 27k cycles
   → Increase to 8S5P = 40 cells, 16.0 Ah × 28.8 V = 460.8 Wh
   → Actual DoD = 119 / 460.8 = 25.8% ✓ (within 25-30% for >50,000 cycles)

Battery mass: 40 × 48 g (18650) = 1.92 kg cells + 0.5 kg structure + 0.3 kg electronics = 2.72 kg
Energy density check: 460.8 Wh / 2.72 kg = 169 Wh/kg ✓ (within Li-ion range)

Step 6: Energy Balance Verification

The fundamental check — does the array recharge the battery during sunlight AND power the loads?

Energy_in (sunlit) = P_sa_EOL × T_sun × η_charge = 1153 × 0.927 × (59.4/60) × 0.90 = 952.5 Wh
Energy_out (sunlit) = P_sunlit_avg × T_sun = 331.7 × (59.4/60) = 328.3 Wh
Energy_out (eclipse) = P_eclipse × T_eclipse = 200 × (35.7/60) = 119.0 Wh
Energy_out (total) = 328.3 + 119.0 = 447.3 Wh

Energy margin = 952.5 − 447.3 = 505.2 Wh → 113% margin ✓ (minimum required: >10%)

Positive energy balance confirmed. Array is conservatively sized.

Step 7: Trade Study

If xlsx skill available → parametric power budget spreadsheet. If thermal skill available → array temperature profile for accurate derating. If orbital-mechanics skill available → beta angle sweep for eclipse/sun variation.

7. OUTPUT TEMPLATE

# [Mission Name] — Electrical Power System Design

## Mission Parameters
| Parameter | Value |
|-----------|-------|
| Orbit | [altitude] km, [inclination]° |
| Mission life | [X] years |
| Orbit period | [X] min |
| Eclipse duration (worst case) | [X] min |
| Sunlit duration (worst case) | [X] min |

## Power Budget
| Subsystem | Safe (W) | Nominal (W) | Peak (W) | Eclipse (W) |
|-----------|---------|-------------|---------|-------------|
| [subsystem] | [X] | [X] | [X] | [X] |
| **TOTAL (with 20% margin)** | **[X]** | **[X]** | **[X]** | **[X]** |

## Solar Array
| Parameter | Value |
|-----------|-------|
| Cell type | [type] |
| BOL efficiency | [X]% |
| EOL efficiency | [X]% (after [Y] years) |
| Array area | [X] m² |
| Array power BOL | [X] W |
| Array power EOL | [X] W |
| Array mass (total) | [X] kg |

## Battery
| Parameter | Value |
|-----------|-------|
| Chemistry | [type] |
| Configuration | [XsYp] |
| Capacity | [X] Ah / [X] Wh |
| DoD (worst case) | [X]% |
| Cycle life at DoD | [X] cycles |
| Battery mass | [X] kg |

## Energy Balance (worst-case orbit)
| Parameter | Value |
|-----------|-------|
| Energy in (sunlit) | [X] Wh |
| Energy out (total orbit) | [X] Wh |
| Margin | [X]% |

## Recommendation
[Selected architecture, key trades, risks, next steps]

8. CLASSIFICATION

9. VARIATIONS

• A: CubeSat (1-12U, <30 W) — Body-mounted Si or GaAs cells, 2S Li-polymer, unregulated bus, peak power tracking, deployable panels for 3U+. Array: 30 cm² per U-face → 7 W/face max. Battery: 20-80 Wh.
• B: LEO Earth Observation (200-2000 W) — Sun-tracking triple-junction wings, 28 V regulated bus, Li-ion 8S packs, 5-7 year life, worst-case eclipse 35 min, 5,400 cycles/year drives DoD < 30%.
• C: GEO Communications (5-25 kW) — Dual wing arrays 50-100 m², fully regulated 50-100 V bus, Li-ion replacing NiH2, 90 eclipses/year (max 72 min at equinox), 15+ year life → EOL degradation dominates. North/south array wings with seasonal tilt.
• D: Deep Space / RTG — Solar arrays impractical beyond ~5 AU (inverse-square: Jupiter gets 50 W/m², Saturn 15 W/m²). MMRTG: 110 W BOL, 73 W at 14 yr. GPHS-RTG: 285 W BOL. Juno exception: 3 huge arrays (60 m²) at Jupiter → 486 W.
• E: High-Power Radar / SAR — Peak power 5-15 kW during imaging passes, duty cycle <25%, requires supercapacitor or high-rate Li-ion for pulse loads, semi-regulated bus with MPPT, array sized for orbit-average not peak.

10. ERRORS & PITFALLS

• E1: Using BOL array power for EOL analysis (triple-junction loses 7-10% over 5 years in LEO — a 1000 W BOL array is only 927 W at 5-year EOL)
• E2: Ignoring temperature derating (cells at 65°C in LEO produce 18% less than at 28°C reference — this is NOT a small correction)
• E3: DoD too aggressive for cycle life (25% DoD @ 50,000 LEO cycles is safe; 40% DoD kills Li-ion batteries in <15,000 cycles = 2.8 years)
• E4: Forgetting battery charge power in sunlit budget (battery charge = P_eclipse × T_eclipse / (T_sun × η_charge) adds 30-60% to sunlit load)
• E5: Body-mounted vs sun-tracking confusion (body-mounted cube averages cos(45°) × 2 faces ≈ 0.47 effective area; sun-tracking gets cos(0°) = 1.0 — factor of 2x difference)
• E6: Ignoring harness losses (5% for small satellites, 3-8% depending on bus voltage and cable runs — at 28 V, 500 W = 17.9 A, and 0.5 Ω harness wastes 160 W)
• E7: Sizing battery only for eclipse, not peak loads (if peak load exceeds array capability during sunlight, battery must cover the deficit — size for max(eclipse energy, peak deficit × duration))
• E8: Using GEO DoD limits for LEO (GEO: 90 cycles/year allows 70-80% DoD; LEO: 5,400 cycles/year demands <30% DoD — same battery, completely different sizing)

11. TIPS

• T1: Start from orbit → eclipse duration → power budget → array → battery → energy balance. Always work in this order.
• T2: Triple-junction GaAs (ZTJ/ITJ) is the default choice for anything beyond CubeSats — 29.5% BOL efficiency at proven cost. Silicon only for lowest-cost missions.
• T3: Rule of thumb for LEO array sizing: P_array_BOL ≈ 2.5 × P_orbit_avg. This accounts for eclipse duty, path losses, degradation, and temperature. Verify with the full formula.
• T4: Always check energy balance, not just power balance. An array that meets peak power but cannot recharge the battery in one sunlit pass will drain to zero.
• T5: Bus voltage selection: 28 V for <3 kW, 50 V for 3-10 kW, 100+ V for >10 kW or electric propulsion. Higher voltage = less harness mass = fewer thermal dissipation problems.
• T6: Margin policy: 20% power margin at subsystem level (PDR), 10% at system level (CDR), 5% at EOL minimum. Never launch with <5% positive margin at worst-case EOL.
• T7: Calibrate against known missions: ISS (110 kW, 2,500 m² arrays, 120 V), Sentinel-6 (900 W, 10.5 m² array, 525 km SSO), Starlink v2 (4 kW, ~30 m² array, 550 km).
• T8: For rapid mass estimation: solar array 3-8 kg/m² (rigid panels with mechanism), Li-ion battery 5-7 Wh/kg at pack level (cells + structure + electronics), PCDU 1.5-3 kg per kW regulated.

12. RELATED SKILLS