Electrical power system (EPS)

Definition

The electrical power system (EPS) generates, stores, regulates and distributes the energy of the spacecraft. In small satellites it almost always follows the same pattern: solar cells generate in sunlight, a rechargeable battery carries the loads through eclipse, and converters and load switches condition and distribute the power. According to NASA’s 2026 technology report, more than 90% of nanosatellites and small satellites use solar arrays with rechargeable batteries[1].

The energy source

At 1 au a surface facing the Sun receives about 1361 W/m²[2]. How much of it becomes electricity depends on the cell efficiency, on the fraction of the surface the cells cover and on the angle of each face to the Sun. Multi-junction III-V cells, the standard choice in small satellites, have nominal efficiencies around 30%; single-junction silicon cells stay near 20%[1].

An uncontrolled spacecraft shows its faces to the Sun in turn and at oblique angles, so the orbit-average power of body-mounted cells is far below what the cell area suggests. For a randomly tumbling convex body the mean projected area is a quarter of the total surface (Cauchy’s theorem), which gives Eq. 1 of the PocketQube review[3]:

with solar constant $S$, cell efficiency $\eta$, packing factor $p$ and eclipse fraction $f_{\mathrm{e}}$.

Architecture

An EPS performs four functions: generation with power-point tracking, storage, distribution and launch safety. Power from the arrays passes through a maximum-power-point tracker (MPPT) or a direct-energy-transfer stage onto the battery bus. Regulators fed from that bus create the fixed voltage rails, and current-limited load switches let the computer turn each consumer on and off and isolate it after a short circuit.

flowchart TB
  subgraph GEN["Generation and tracking"]
    direction LR
    SA["Solar arrays"] --> MPPT["MPPT converter"]
  end
  subgraph STO["Storage"]
    direction LR
    BCR["Charge control, cell balancing"] --> BAT["Li-ion battery and heater"]
  end
  subgraph DIST["Distribution"]
    direction LR
    REG["5 V and 3.3 V regulators"] --> LS["Current-limited load switches"]
    LS --> LOADS["OBC, COMMS, ADCS, payload"]
  end
  MPPT --> BCR
  BAT --> INH["Separation switches and RBF pin"]
  INH --> REG

Solar cells

The specific power of the array products in the report clusters around 30 W/kg[1]. For deployable arrays, the mass of hinges and release mechanisms must be counted with the panels.

Body-mounted power

With $\eta =0.28$, $p=0.6$ and $f_{\mathrm{e}}=0.36$ for a 525 km Sun-synchronous orbit, Eq. 1 gives the orbit-average power of fully loaded spacecraft:

The ratio of about one quarter between a 1P and a 1U matches the scaling reported by Bouwmeester et al., and the 3P value agrees with the roughly 1 W of Delfi-PQ, which carries cells on its long faces only[4]. Per kilogram a PocketQube collects more power than a CubeSat; what it lacks is absolute power for a single consumer.

Eclipse

The eclipse fraction depends on altitude and on the beta angle between the orbit plane and the Sun direction. For a cylindrical shadow:

Dawn–dusk Sun-synchronous orbits sit at high beta and may avoid eclipse for most of the year; noon–midnight orbits see the longest eclipses. The EPS must be sized for the worst case of the mission, not the average.

Battery

The report gives operating ranges of 0 to 45 °C for charging and −20 to 60 °C for discharging commercial Li-ion cells[1].

Distribution and bus voltages

Traditional small-satellite EPSs use a main battery bus of 8.2 V, two Li-ion cells in series (2S), and derive regulated 5.0 V and 3.3 V outputs from it; CubeSats often run on a single low-voltage bus. Gallium-nitride converters make higher buses of 12, 28 or 50 V practical, and the power handled by the distribution systems of new high-power small satellites has more than doubled[1].

Power budget

A power budget works backwards from the load profile to the power the arrays must deliver in sunlight. Eclipse loads are supplied by the battery and repaid with charging losses, so they enter with a lower path efficiency[5]:

where $P_e$ and $P_d$ are the mean loads in eclipse and in daylight, $T_e$ and $T_d$ the corresponding durations, and $X_e$ and $X_d$ the efficiencies of the paths through the battery and directly to the loads. The reference 1P below carries a mean load of 134 mW, which leaves an energy margin of 284% per orbit and a depth of discharge of 4.3% in eclipse.

The power budget article builds this budget step by step and shows how the transmitter duty cycle drives it.

Eclipse and cycle life

At 525 km with the Sun in the orbit plane, 37.5% of every orbit is spent in shadow, which means about 5500 charge and discharge cycles a year. Battery ageing is driven by depth of discharge, temperature, charge and discharge rates and storage conditions[1]. For multi-year missions, keeping the depth of discharge low usually extends life more than adding capacity.

Cold charging

Charging lithium cells below about 0 °C plates lithium on the anode and permanently reduces capacity. Small satellites have little thermal mass, and their batteries cool quickly in eclipse: a 250 g 1P falls to about −11 °C in a single-node model at 550 km[6]. For SMOG-1 the conservative thermal model predicted the battery below 0 °C in eclipse, so charging was restricted to about half of the sunlit period[7]. A temperature threshold in the charge controller and a small heater are needed even at the 1P scale.

Cells and boards

Body-mounted triple-junction cells are grown on germanium, which expands by about 6 ppm/K, whereas FR-4 panels expand by 14–17 ppm/K in plane. The mismatch cracks cells and fatigues solder joints under thousands of thermal cycles[8]. Screening with the Engelmaier model gives a median life of about 5100 cycles, less than a year, for a rigid joint on a 1P that swings by 30 K per orbit[6]:

Compliant attachment, smaller cells and CTE-tailored substrates move the joints back into a safe regime.

Launch safety

The CubeSat Design Specification requires the power system to be off from delivery to the deployer, at least one separation switch that electrically isolates the power system from the powered functions while in the deployer, and batteries with charge and discharge protection against unbalanced cell states[9]. Battery chemistry and capacity also decide which extra safety documents and tests the launch provider requires.

Ageing and end of life

Radiation lowers the efficiency of the cells over the mission and cycling lowers the capacity of the battery. A power budget must therefore close with end-of-life values, not with beginning-of-life ones.

At PocketQube scale

The binding constraint of a 1P is not energy per kilogram but power for a single consumer. A transmitter amplifier that draws 1.25 W while sending 0.5 W of RF is more than twice the orbit-average generation; it can only run at a low duty cycle, with the battery bridging the bursts. Grizu-263A filed an average generation of 0.75 W and an active consumption of 1.5 W for its frequency coordination, so its high-consumption mode was necessarily duty-cycled[3]. A deployable Sun-pointed array can lift a 3P above 10 W[4], at the cost of attitude control and mechanism risk.