How to Size Solar Arrays for Your Spacecraft Mission

The power system is often called the "heartbeat" of a spacecraft: every other subsystem depends on it. Solar arrays are the primary power source for virtually all Earth-orbiting satellites and many interplanetary missions, and their sizing drives mass, cost, and configuration in ways that cascade through the entire design. This guide presents the engineering methodology that aerospace power system engineers use to size solar arrays, from first principles through practical implementation.

Step 1: Establish Your Power Requirements

Before you can size arrays, you need a complete power budget. This means accounting for all operating modes:

• Sunlit nominal operations: all subsystems running at typical power draw
• Eclipse operations: subsystems that remain on during eclipse, plus battery charging requirements
• Peak load events: deployments, maneuvers, high-power transmissions
• Safe mode: minimum power to maintain spacecraft health

Power budgets should include margin: typically 20–30% at the component level and a system-level margin of 10–20% on top. Margins are not padding; they account for measurement uncertainty, component variation, and degradation that is difficult to predict precisely.

Step 2: Determine Solar Flux and Pointing Geometry

The solar constant at 1 AU (Earth's mean distance from the Sun) is approximately 1,361 W/m². For missions at other distances, flux scales with the inverse square of distance: a spacecraft at 1.5 AU (roughly Mars) receives about 590 W/m².

The effective power density on your array depends on:

• Solar incidence angle (θ): power scales with cos(θ); a body-mounted array with imperfect sun pointing loses efficiency rapidly as angle increases
• Seasonal variation: Earth's orbital eccentricity means solar flux varies ±3.4% between aphelion and perihelion
• Array orientation strategy: single-axis vs. two-axis gimbaling, body-mounted fixed panels, or deployable wings all impose different constraints

Step 3: Account for Eclipse Fraction

In low Earth orbit, satellites spend a significant fraction of each orbit in Earth's shadow. For a 550 km circular orbit (roughly Starlink altitude), the orbital period is approximately 95.5 minutes and the maximum eclipse duration is about 35 minutes, giving a worst-case eclipse fraction near 37%. Sun-synchronous orbits at some local solar times can achieve near-continuous sun exposure.

The eclipse fraction directly determines battery sizing and affects how much power the array must generate during sunlit periods to both support loads and recharge batteries. The energy balance equation is:

P_array × t_sun × η_harness = P_load × T_orbit + P_batt_charge × t_eclipse

Where η_harness accounts for wiring and power conditioning losses (typically 0.85–0.92 combined).

Step 4: Select Solar Cell Technology

Solar cell selection is a key cost-performance tradeoff. The main options are:

• Triple-junction GaAs (3J): Industry standard for commercial and government satellites. Beginning-of-life (BOL) efficiency of 29–32% for production cells; heritage from Azur Space, Spectrolab, SolAero. High cost (~$300–600/W for array-level systems) but excellent radiation tolerance.
• Inverted metamorphic (IMM) and four-junction (4J) cells: Up to 33–35% BOL efficiency; used in high-performance missions where mass is the dominant constraint. Higher cost than 3J.
• Silicon solar cells: ~14–16% efficiency; rarely used in modern spacecraft due to poor radiation tolerance and lower efficiency, but very low cost
• Perovskite and advanced III-V: Lab efficiencies exceeding 40% for multijunction designs; not yet space-qualified but an active area of development

Step 5: Apply End-of-Life (EOL) Degradation Factors

Solar arrays degrade in the space environment primarily due to particle radiation (protons and electrons trapped in the Van Allen belts, solar energetic particles) and ultraviolet exposure causing coverglass darkening and adhesive degradation. For mission sizing, you must design to end-of-life power requirements, not beginning-of-life.

Key degradation factors to apply to your BOL power:

• Cell degradation (radiation): Typically 2–8% per year in LEO depending on altitude and inclination; use SPENVIS or similar tools for orbit-specific fluence analysis
• Coverglass transmission: ~1–2% total mission degradation for standard ceria-doped covers
• Thermal cycling fatigue: Interconnect cracking in large arrays; mitigated by design margins and interconnect flexibility
• Array temperature: GaAs cells have a temperature coefficient of approximately −0.20% per °C; arrays in GEO may reach 60–80°C in sun, reducing output 5–10% from STC

A combined EOL efficiency factor (the product of all degradation multipliers) of 0.75–0.85 over a 15-year GEO mission is a reasonable starting range for preliminary sizing.

Step 6: Size the Array

With all factors established, the required array area is:

A = P_EOL_required / (S × cos(θ) × η_cell × η_EOL × η_harness)

Where S is the solar flux (W/m²) and η values are the respective efficiency and degradation factors. For a preliminary design, iterate between array area, mass (using typical areal density of 1–3 kg/m² for rigid panel arrays, 0.5–1.5 kg/m² for flexible deployables), and power conditioning architecture until the system closes.

Common Sizing Mistakes

• Using BOL efficiency when EOL is the correct design point
• Neglecting array temperature: GaAs performance degrades significantly above 25°C
• Underestimating harness and PCDU losses, particularly in distributed architectures
• Ignoring worst-case sun angle in the mission timeline (e.g., winter solstice for sun-synchronous, worst beta angle for LEO)
• Treating power margins as optional; they are not

For orbital parameters relevant to your mission's eclipse fraction and sun angle analysis, use the SpaceNexus Orbital Calculator.