1. The panels that must keep working

Andy Weir’s The Martian returns as the recalled work because Mark Watney’s power budget is simple and brutal: solar panels during the day, batteries at night, and every watt accounted for because there is no utility company to call. A LEO desktop has the same structure, just with more eclipses and more customers arguing over the watts.

This entry reads about how spacecraft generate, store, and regulate power.

2. The electrical power subsystem

An Orbital Radar overview defines the electrical power subsystem (EPS) as the part of a spacecraft that generates, stores, regulates, and distributes all electrical power. In LEO the standard architecture is:

  • Solar arrays convert sunlight to electricity during sunlit periods.
  • Rechargeable batteries take over during eclipse and handle peak loads.
  • Power conditioning and distribution electronics hold the main bus at a steady voltage and protect each circuit.
  • Loads draw from the bus according to their operating modes.

A Power Electronics News article on CubeSat and SmallSat power architecture describes how these systems balance generation, battery management, and size-weight-power-cost constraints.

3. Solar arrays

Solar arrays are the default primary power source in LEO. Key considerations include:

  • Cell technology: triple-junction cells currently offer the best trade-off between efficiency, mass, and radiation degradation.
  • Body-mounted vs. deployed: body-mounted panels are simpler and more robust; deployed arrays produce more power but add deployment risk.
  • Sun-pointing: arrays that track the Sun produce more energy but need pointing mechanisms and clear fields of view.
  • Degradation: radiation, ultraviolet, and atomic oxygen reduce output over the mission life; arrays are sized for end-of-life performance, not beginning-of-life.

4. Batteries

Batteries are the spacecraft’s night shift. A satellite battery primer notes that LEO satellites experience roughly 5,840 eclipses per year, each 30 to 36 minutes long. Modern spacecraft use lithium-ion batteries because their energy density is roughly three times that of the older nickel-hydrogen chemistry.

Battery sizing is governed by:

  • Eclipse energy requirement: how much energy the spacecraft must draw while in shadow.
  • Depth of discharge (DOD): how much of the battery’s capacity is used per cycle; lower DOD means longer cycle life.
  • Cycle life: how many charge-discharge cycles the battery can survive at the chosen DOD.
  • Temperature: batteries work best within a narrow temperature band and may need heaters or thermal control.

5. Power conditioning and distribution

The PCDU or EPS controls the interface between arrays, batteries, and loads. Functions include:

  • Maximum power point tracking (MPPT): adjusting the array operating point to extract maximum power as illumination and temperature change.
  • Battery charge management: regulating charge current and voltage to avoid overcharge and to balance cells.
  • Bus regulation: maintaining a stable voltage for the rest of the spacecraft despite varying generation and load.
  • Fault protection: current limiting, latching switches, and overvoltage protection to isolate faults.

A 3Cat-8 mission paper describes a typical small-satellite EPS with array conditioning, power distribution, and a battery pack sized to subsystem power consumption.

6. What this changes

  • The desktop’s power system will almost certainly be solar-array-plus-battery in LEO.
  • Solar array sizing must account for eclipse, degradation, pointing, and customer payload growth.
  • Battery sizing must account for cycle count, DOD, and thermal control.
  • The next entry will read about power budgets and eclipse energy balance.