Orbital Relay for Communications Access (ORCA)
ORCA is a proposal for a Lunar Communications Constellation that seeks to implement part of the NASA LunaNet concept. The spacecraft will provide uplink and downlink to surface and orbiting assets with intermittent and occasional real-time relay to Earth. The mission brief called for a commercial approach to costing and scheduling, requiring at least 3 satellites to be launched as payloads on one or more ESPA rings. My role on this project team was thermal lead, which is also the focus of the work below. I also designed graphics and images.
This work was completed for a group term project in ASEN 5148 - Spacecraft Design with Dr. Daniel Kubitschek and Dr. Christopher Grasso at CU Boulder.
ORCA Thermal System Design
ORCA orbit configuration
Lunar IR flux with the sun in the +X direction
Diagrams of a single inclined satellite in two beta angle conditions (left) and the Moon-Earth-Sun system from the perspective of the Moon showing the effect of changing beta angles over the year (right)
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.
Thermal System Design Overview
The guiding design concept to respond to this highly variable environment and mission needs is to maximize the thermal stability of the spacecraft: increase thermal mass, maximize internal conduction paths, and insulate the bus to close the gap between the hot and cold cases and reduce the satellites reactiveness to its changing environment.
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 shown for PDR, assuming a right-sized spherical spacecraft with averaged surface properties. 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.
Heater Allocation
References
Ababneh, M. T., et al. “Copper-Water and Hybrid Aluminum-Ammonia Heat Pipes for Spacecraft Thermal Control Applications.” In International Heat Pipe Symposium (IHPS), June 2018, (No. JSC-E- DAA-TN54602).
Baker, Charles, et al. ”Lunar Reconnaissance Orbiter (LRO) rapid thermal design development.” Heat-pipes for Space Applications International Conference. 2009.
Beckman, Mark. ”Mission design for the lunar reconnaissance orbiter.” 29th Annual AAS Guidance and Control Conference. No. AAS-07–057. 2007.