I’m giving it all she’s got, Captain!
–Montgomery “Scotty” Scott, Star Trek
This post is part of a series of updates on our mission planning process, which began with Mission Planning Updates, Part 1: Trajectory.
Solar energy is abundant in the inner Solar System. Therefore, solar panels are a ubiquitous choice for powering spacecraft. For shorter missions, such as Space Shuttle flights or launch vehicle upper stages like Centaur, energy storage such as fuel cells or batteries may be all that is needed. But for lengthy missions, spacecraft need to generate their own power to sustain mission operations.
Apart from those few that use radioisotope thermoelectric generators (mostly deep space probes such as Voyager and Pioneer), that is done with photovoltaics. Our own spacecraft use surface-mounted Solaero ZTJM triple-junction cells, together with a GomSpace P31u for power management, including a lithium-ion battery for energy storage.
Our spacecraft, like many, will not produce enough power to continuously run all of its components at once. Their surface-mounted solar panels produce at most 10 W for each spacecraft, while the electrolyzers alone consume about 6 W, for example. High performance lithium-ion batteries mean we will be able to run a power deficit for many hours, but not forever. Therefore, we will need to pick and choose carefully how to match the available power to the needs to the mission at all times.
Planning a power budget for the spacecraft cannot be totally decoupled from other mission planning elements, such as the trajectory previously discussed, and the thermal concerns that will be the subject of the next post. We have already developed detailed power budgets and analysis for the spacecraft through their entire trajectory, but the scripts we made to do this will need to be re-run on the new trajectory data after it is finalized.
In the rest of this post, I want to explain a bit about why the spacecraft trajectory and attitude are linked to the power budget, and then discuss the results of our existing power analysis and how we expect it to change this semester.
As our spacecraft traverse their planned path as described in the first post of this series, one might expect the solar power available to it to change. For example, the solar irradiance at Mars is about 44% that at Earth. Although the distance from the Sun does affect the solar irradiance, it has an inverse-square relationship with distance from our star:
This is because the light emitted from the Sun spreads out like an expanding sphere as it moves away, and the surface area of a sphere increases proportionally to the square of its radius. However, this is not a very significant effect in our case.
Our spacecraft will be deployed close to the Earth and remain in its orbit (or the Moon’s) for the entire mission duration, meaning they will never be more than about 1 million km closer or father from the Sun than the Earth itself is.
Because the Earth is about 150 milion km from the Sun already, that means the solar irradiance on the spacecraft will vary on the order of 1%. And our spacecraft will spend most of their time closer to the Moon’s distance or in its orbit, which is even closer to the Earth, and so solar irradiance will vary even less there.
There is one important effect the trajectory has on our power generation, however: Occasionally, our spacecraft will enter an eclipse where the Sun is blocked by the Earth or Moon. In these cases, there will not be any power available from the solar panels at all. The spacecraft will be dependent on its battery reserves. Eclipses also pose a danger to the water in the propellant tank, which is susceptible to freezing, but that is a topic for another post.
Our spacecraft use surface-mounted solar panels, which have the advantage of being mechanically simpler than deployable solar arrays such as those on the ISS, but have a limited surface area in comparison, and also no ability to orient the solar panels towards the sun independently of the spacecraft.
Because of this, the power generated by the spacecraft is dependent on orientation. All orientations produce some power, but the amount varies around an average of about 7 W. Our spacecraft have an excess propellant budget for their attitude control cold gas thrusters, so it may be advantageous to reorient the spacecraft to face the sun when power is needed for long durations. At other times, it may not be worth the propellant cost to do so, if the spacecraft will not need to perform high power operations, or if it does but there is time to wait for the batteries to recharge in between them.
The above images are the results of the previous power analysis. Our next run will show improved performance, but more on that later. The first script takes as an input the spacecraft solar panel layout to determine power production as a function of spin axis angle relative to the Sun.
The next uses the result of the first, plus the trajectory and orientation of the spacecraft as inputs, as well as the ephimerides of the Sun, Earth, and Moon. With all of this information, we can compute the power production of the spacecraft across the planned mission.
The final step is to add in the planned power consumption across the trajectory. This means determining how much power each component uses, when it will be needed, and at what duty cycles. By subtracting production from consumption over the entire trajectory, we can then compute the net power production of the spacecraft across the whole mission.
We can then assess the times when net power production is negative against the total energy capacity of the cells in the P31-u power unit. Even in this outdated analysis, we found that the spacecraft would never go below 40% depth of discharge (60% battery capacity) in the nominal mission.
It is possible to alter the mission planning to improve this, by for example, setting a DOD limit and halting electrolysis at this limit to allow the battery time to recharge. In our case, the final design of the spacecraft will improve the results as well.
We expect the outlook to be more favorable in our upcoming power analysis this semester, for two main reasons. First, the final spacecraft design has a greater solar panel surface area, especially on the spinning edge sides. Therefore, the spacecraft produces more power at all orientations than it did previously. The improved power production means the spacecraft will nearly always produce more power than the electrolyzers for the main thruster consume, which was previously not the case. This means the spacecraft will be able to continuously thrust in most orientations, instead of needing to take breaks every few hours to recharge. This eliminates several of the negative net power cases in the existing analysis.
Second, with our improved ground station and other changes to the communications architecture, we were able to reduce the required power draw of the spacecraft transmitter. This was previously the highest power draw item on the spacecraft (that honor now belongs to the cold gas thruster solenoid valve) and will be needed frequently during one of the most critical phases of the mission, because the spacecraft will be downlinking as much as several minutes every hour when working to verify lunar orbit. Therefore, the spacecraft consumes less power than it did previously.
Because the spacecraft now consumes less power and produces more power than it did previously, the expected depth of discharge will be decreased (improved) and the spacecraft will spend more time with its batteries at a high state of charge. This will ensure it is prepared for emergencies and unexpected net power drains, for example, if a solar panel were to fail and the spacecraft needed to reorient to begin producing power again.