This post is part of a series of updates on our mission planning process, which began with Mission Planning Updates, Part 1: Trajectory. We now turn to considering the effects of thermal radiation on our spacecraft, and the analyses we have performed and will repeat to ensure mission success.
We discussed solar radiation in our previous post, but only visual light for photovoltaics powering our spacecraft. As everyone is familiar with, sunlight also imparts heat.
Is space hot or cold? The answer is a confusing one, because the vacuum of space is an excellent insulator. Convection and conduction are nonexistent (outside of the spacecraft itself) with almost nothing to carry heat to or from the spacecraft. The only significant form of heat transfer to and from the spacecraft is radiation.
Keeping the spacecraft at a safe and stable temperature requires balancing the incident thermal radiation (mainly from the Sun) with the emitted thermal radiation from the spacecraft, as well as heat generated onboard. Larger, more complex spacecraft have dedicated radiators they can deploy for this purpose, but our CubeSats do not. In our case, balancing the radiation on the spacecraft means reorienting it to adjust the amount of surface area that is exposed to the Sun, and carefully managing the heat generated onboard.
Our spacecraft have two thermal goals to keep in mind. It is important that the water in the propellant tank remain liquid, because even though there is enough room in the tank for it to expand and freeze, it cannot be used as propellant until it thaws! Additionally, a frozen propellant tank would eliminate the slosh damping effect, harming our attitude control strategy.
In addition, it is vital that the spacecraft electronics, particularly the batteries, are keep cool to prevent potentially destructive overheating. This goal may seem to be add odds with the need to keep the water heated, but the two are actually complementary. By heat-sinking the spacecraft electronics to the propellant tank, we are able to keep the water warm at the same time as we remove waste heat to keep the electronics from overheating. This allows for a symbiotic relationship between subsystems, a design philosophy we have made use of in our architecture wherever possible.
However, in space, that heat needs to go somewhere eventually, and this is where the spacecraft orientation comes in. In the previous post, we discussed how the spacecraft naturally receives the most solar radiation when its spin axis is oriented towards the Sun, exposing the largest of its surface area.
Passive Thermal Analysis:
In our analysis we mainly consider two extremes:
- Can the spacecraft keep its propellant tank from freezing indefinitely in its minimum solar incidence orientation? Also, can this be done while maintaining a positive power balance (still being able to charge the batteries, instead of needing to spend more power heating the spacecraft than it receives from solar panels in this orientation)?
- Can the spacecraft keep its electronics (primarily, the batteries) from overheating in its maximum solar incidence orientation? Also, does being in the maximum solar incidence orientation retain all available mission operations, or will some be precluded by fear of overheating?
In our design process, we have considered these two questions repeatedly using finite-element analysis, iterating on the structural design of the spacecraft, the locaion of the electronics within it, and the thermal coating applied to the surface, until the answer to both questions was “yes.”
Our modeling included both steady state and transient analyses, with the latter including cases such as the spacecraft operating its radio amplifier or firing its thruster (producing lots of heat) in the maximum solar orientation, or crashing and losing its ability to heat itself temporarily during the minimum solar orientation.
The end result of all of this is that we were able to refine our thermal design until we finalized it last year, with:
- A margin of 10 C above freezing while maintaining net power production, and a margin of 5 C above freezing while running at minimum power, in minimum solar orientation.
- A margin of 15 C below dangerous temperatures while activating the spacecraft radio in the maximum solar incidence orientation. We also found that the thruster pulses do not take place often enough to significantly heat the spacecraft.
Thermal Mission Planning:
The steady-state thermal analysis above might give the impression we do not need to factor thermal concerns into our mission planning. Thanks to our careful structural and coating design, the spacecraft will passively remain at acceptable temperatures even during maximum power output in maximum solar incidence, or during minimum power output in minimum solar incidence.
However, there are still a few considerations to be made. Most notably are eclipses, when the spacecraft position is such that the Sun is obscured by the Earth or Moon. This does not happen too often or too long in our trajectory, but can be dangerous. The spacecraft receives no power and very little heating in this situation, meaning it must rely on cranking up onboard heaters to survive. But without solar power, those heaters are sustained by the batteries, giving the spacecraft a finite time until it can no longer keep the water liquid.
The spacecraft are able to survive an eclipse of any duration that might be encountered in our planned mission, but we must be careful not to let them be caught by surprise. A spacecraft that has just finished massively discharging its batteries to perform some sequence of maneuvers may not be prepared to keep itself heated for as long as might be necessary.
This is part of the motivation for the power and trajectory analysis in the previous posts: ensuring that the spacecraft enters any eclipse in a state of maximum charge to be best prepared to survive until it can see the Sun again.