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Friday, 25 July 2014

My First Month at NASA, Part II: Passive Thermal Design

Hello again! Last time, I updated you on the progress I have been making at the Johnson Space Center, NASA's center for human spaceflight! We had looked at some of the ideas for configuration and deployment and had been shown animations. This time, we're going to get more technical, and I'm going to show you what conclusions I've been making, and how I made them, concerning the thermal control system for the conceptual spacecraft!

If you've been following along, you'll recall that I'm working on a spacecraft which features active radiation shielding. Passive radiation shielding involves the use of material to simply block incoming radiation, and is the technique which is used today. Outside of Earth's protective magnetic field, many studies suggest that passive shielding may not be enough, and with many groups becoming more interested in missions to Mars, radiation protection becomes that much more important. Active radiation shielding involves a spacecraft generating its own magnetic field to block and deflect incoming radiation.

The concept being developed in NASA's Innovative Advanced Concepts study (NIAC), involves using large coils of superconducting wire to generate large magnetic fields around a spacecraft habitat. The Phase I study can be found here. My internship involves building on this concept and helping to wrap up the Phase II study.

In order for these coils to work, to be superconducting and thus more capable of generating and sustaining the large magnetic fields, the coils must be kept cold. How cold? Well, we have estimated that the coils must be kept below 40 Kelvin (-233°C, -387°F) in order to work.

You might be tempted to think, "Well, they'll be working in the cold vacuum of space, for deep space missions, that shouldn't be a problem, right?" While it is true that deep space is cold (~3 K), it is important to remember that this system will be launched from Earth and that the Sun outputs ~1367 Watts/square metre at Earth's orbit. Considering the system is envisioned to have 6 cylindrical coils, with a length of 20 metres and a radius of 4 metres, we're looking at a surface area of 603 m/coil, amounting to 824 kW of heat on the surface of 1 coil!

For my internship here at NASA, I have been assigned the task of designing the thermal control system for the spacecraft, concerning myself mostly with keeping the coils below 40 Kelvin. So, the first thing I did was refresh myself on the system and the working environment.

The coils will surround a habitat, similar to NASA's Destiny module on the ISS. They will have dimensions as mentioned above, and the concept is that they will begin deployment in high-Earth orbit (HEO). My first thought concerning the working environment was that there would several sources of heat including: the Sun, the Earth, the Moon, and the spacecraft itself.

After reviewing the literature, I decided that the infrared radiation from the Earth and Moon would be almost negligible in HEO. As can be seen in the above links, the heat observed at geosynchronous orbit (GEO, closer to the Earth than HEO), was approximately 5 Watts/m^2 each. This is small compared to the 1367 Watts/m^2 coming from the Sun. I deduced that I could incorporate the Earth and Moon's effects into an appropriately sized margin.

With the Sun being the dominant heat source, I began investigating ways to block it, and radiate the most heat away from the spacecraft. The first thought I had was the use of a sunshield. The James Webb Space Telescope will be using 5 layers of special material (aluminized kapton) in order to provide a cool side of 40 Kelvin. How perfectly convenient! A video of its deployment can be found here.

Looking back at the system, I remembered that the spacecraft will require solar panels, and that these panels had not been really sized yet. By this, I mean that while it was understood that they would be needed, no one had set size requirements, meaning that I was allowed to move them and change their size. Moving them to one side of the spacecraft, I realized I could block the majority of the sunlight, and use a heat shield to block the rest, as well as block the heat radiating from the solar panels.

This much you might be familiar with, if you have been reading my earlier posts. Now, to go into the calculations.

As with many calculations, the key is all about balance. The spacecraft must radiate out as much, or more, of the heat coming in. I have outlined some of the calculations below.

In the top, ideal, case, the thermal load on a spacecraft, is determined by the heat coming in from the Sun, Earth, Moon, etc., and the internal heat load.

Qexternal + Qinternal < Qradiating

In our case, the Moon and Earth are not making much an effect, so the equation for the external heat load is:

Qext = Area (of the coils),*alpha*Solar flux

where alpha is the absorptivity of the coils and the solar flux is ~ 1367 Watts/m^2. The absorptivity is a value attributed to an object's ability to absorb solar ultraviolet radiation.

So, in the ideal example, you would want to make the area and the absorptivity low. You would want the smallest area to be exposed to the Sun, and we would want the area which is exposed to absorb as little thermal radiation as possible.

The internal heat is determined by the heat from the electronics, any heating which is required (to keep the astronauts comfortable, for example), and any other sources. In the Phase I study, it was determined that this thermal load would be approximately 390 Watts. This heat must be released from the system or the system will heat up, rendering the coils useless.

The radiating heat is determined by the area of the radiator, its temperature, and emissivity. The emissivity is the value attributed to a surface's ability to reflect infrared radiation. Generally, white and reflective surfaces have a high emissivity, and black surfaces have a low emissivity. The dependence on temperature carries some interesting consequences. For example, a space radiator starting at 300 K (roughly room temperature) radiates more heat from its system than a cooler radiator. And since the relationship is to the power of 4, this dependence becomes very important.

Now, in our system, I decided to block much of the incoming solar radiation with solar panels, and then block the heat radiating from the back of those panels using a heat shield. This is depicted in the lower part of the picture. The heat shield is the tall rectangle, and the larger rectangle is how the cylindrical coils may appear from a distance.

At first, I thought I had to trace the heat coming from the Sun, off the back of the panels, blocked by the heat shield, and radiating onto the coils. I discussed this with a NASA colleague, who helped me with the equation featured above: Q = F x A x e* x sigma x (T1^4 - T2^4). Do you see why I chose to take a photo of the equation? Easier to represent when I can write it down.

Let us look at this more closely. The F is the view factor from the panels to the heat shield, thus labeled F12. View factors are simple geometrical equations used to calculate how much of one surface another can see. For example: two plates of equal size, facing each other, can "see" each other perfectly. If heat were coming off of one plate, it would be able to be transmitted easily to the other, and its view factor would be 1. If these same two plates were perpendicular to each other, the radiating heat would have a difficult time making it to the other plate (as the radiation travels in a straight line), so its view factor would be 0. In our case, the view factor from the solar panels to the heat shield is assumed to be 1.

'A' is the area of the solar panels. Next, the e* variable. The equation used above combines the process of going from the back of the panels to the coils in a single equation, and the e* variable is what allows us to do this. The e* is still emissivity, but now it is the emissivity for the heat shield. Most space-worthy insulating materials have an emissivity around 0.01, and the James Webb Space Telescope's heat shield will have an emissivity of 0.000001. Therefore, it is assumed that our heat shield would be somewhere in between those two values, depending on the robustness of our design.

Finally, the temperatures assumed are 373 K for the back of the panels (a common value, as indicated in several papers, and by my NASA colleague), and 30 K for the coils themselves.

Now, these equations are simple enough, but understanding the complexities of them, well, that has taken quite a lot of my time. I spent a long time reviewing thermal engineering design and the concepts involved, so I could better understand these equations, which area to use in which equation, so I could truly understand the system. By doing so, I have a much better understanding of what is involved in this system, and what minute things I could change in order to design the system to fit my needs and requirements.

This understanding lies at the root of engineering. Once you understand the system, you can begin to manipulate the options. In one satellite design assignment, my team was finding that the mass of our satellite was much too high, surpassing our expectations by 2 000 kg. We had been trying to change factors which would largely and blatantly affect the mass of the system. However, our professor, with his years of experience, went into the design of the ground system, changed one small variable concerning the communication, and was able to improve the efficiency of the ground system, which reduced the size of the power system required on the spacecraft, which reduced the size of the onboard batteries, thereby reducing the propellant needed to sustain the satellite and dropping the total satellite mass by 2 000 kg. These kinds of "tricks" are what make an excellent engineer, someone who understands the complexities and inter-connections and can understand ways to change and design the system more effectively.

At this time, I am very convinced of my calculations regarding the passive thermal control of our spacecraft. I have different options for the heat shield, some of which can thoroughly block and radiate the heat. Some options require the use of active thermal control, such as onboard cryocoolers, and this begins my next step; researching these systems.

As I mentioned before, the technique is all about balance. Do I recommend a large, very efficient heat shield, and no onboard coolers? This would certainly eliminate the power requirements of onboard systems, and the mass therein, but there would be added mass and complexity due to the use/fabrication of the heat shield. Plus, there would be no redundancy. Could I get away without using a heat shield? I do not think so. The thermal load coming from the Sun seems to greatly surpass current space-ready cooling systems, so I believe something is required to decrease the strain on the system. Perhaps, I should use a "medium-level" designed shield and onboard cryocoolers. If so, what systems will I use, and what will their mass and power requirements be? These questions and more I am looking to answer in the next month.

It has been a busy time here in Houston. I have been trying to get out more and see things, but I have also not been as inspired to do so. The city of Houston does not call to me as other cities have, and the security requirements on-site at NASA have limited my exploration. I did finally take the NASA center public tour, which you'll see in a later post!

In my down time, I am working on my IAC paper (refining my work on galactic cosmic radiation to present in Toronto this September), final presentation/report for the ISU, finding and securing a PhD, and trying to enjoy any spare time that I have. It has been a difficult balance, but I'm working on it.

As always, thanks for reading, live long and prosper! 

1 comment:

  1. I love your blogs, especially this one. Keep the updates from Houston coming!