Lunar Thermal Environment

ORCA consists of 3 satellites orbiting the Moon in circular orbits: one is in an equatorial orbit and the other two are in offset (RAAN and AOP) inclined orbits at i=70°.

In lunar orbit, the ORCA satellites experience incident solar gain similar to that at Earth, minimal albedo, and surface IR flux ranging from minor to substantial. These conditions result from the low conduction of lunar regolith, the lack of atmosphere to convectively distribute heat, and slow rotation period of the Moon, creating a hot spot on the sun-facing side.

These conditions create a dynamic thermal loading environment that drives design parameters, as well as a significant gap between hot and cold loading cases. Two key orbit parameters determine the cases: beta angle and inclination.

Design Cases. ORCA’s thermal system is designed for bounding hot and cold cases. The hot case occurs when beta angle = 0°, when the spacecraft passes between the surface IR hot spot and the sun on every orbit. This is always true for the equatorial satellite.

The cold case occurs when beta angle = 90°, when the spacecraft are in full sun but always miss the IR hot spot. This causes the inclined satellites to be the driving case.

The system block diagram shows thermal system elements and their interfaces with the environment and key subsystems, including propulsion, mechanisms, reaction wheels, batteries, CDH.

A single-node analysis is likely sufficient for PDR, assuming a right-sized spherical spacecraft with averaged surface properties and no pointing. This analysis shows the initial temperature difference between the hot and cold cases and that a relatively passive design can somewhat close that gap. However, fully closing the gap requires ~250 W of heaters in the cold case and a cold-biased spacecraft. Designing for a warmer operating temperature compromises end-of-life (EOL) performance, after surface coatings have degraded due to UV exposure, an effect that can have an impact as early as 1,000 hours into the mission.

Due to the particular IR loading conditions for this mission, the analysis was taken slightly further to consider the right approximate shape of the spacecraft along with direction-specific surface treatments. The hot and cold cases are significantly closer at beginning of life with minimal heater power. At the end of life, the spacecraft is still approaching the upper operating temperature but no longer requires heaters. The multi-node analysis also adds a survival case, in which the lower temperature bound is allowed to reach -20°C, temporarily compromising the propulsion system.

Component Selection

Emphasis on a short development schedule and heritage technology led to the selection of COTS components wherever possible. Some key selections were passive radiative louvers, constant-conductance heat pipes, and significant MLI coverage.

Passive Components

• Multi-Layer Insulation (MLI)

  • 5 sides of the spacecraft; a=0.14, e*=0.05

• Radiative Coating: 6th side; Silver Teflon: e = 0.87

• RTD Thermostat

  • At least one at major subsystem components

• Thermistors

  • Accompany each heater & other components

• Radiative Louver

  • Closed: e* = 0.14; Open: e* = 0.74

  • Flying 2, 0.32 m2 radiating area

Active Components

• Heaters: Flexible polyimide thermofoil

  • Specifying 12 in 3 sizes at this stage:

    • S: 25.4 x 25.4 cm, 8W

    • M: 25.4 x 76.2 cm, 24W

    • L: 50.8 x 76.2 cm, 53W

• Heat Pipes

  • Constant Conductance, Aluminum-ammonia

  • Expected 15-64 W heat transfer capacity, dependent on gradient

Heater Allocation