At the last IAC, Elon Musk presented the latest changes to the BFR.
Throughout this post, I will be referring to what was previously the ITS as the BFR, as that appears to be what SpaceX now refers to it as externally now. The interplanetary spaceship will be refered to as the BFS, and the booster will be called the BFB.
Some of the numbers were changed between the presentation and the PDF being posted. One of these changes was a factual error, about the internal cabin volume of the BFS, and the other was a change of the total BFB thrust, and the removal of the total BFR mass. For this post, I will mostly use the numbers from the presentation, as they are more complete, but I will note where the numbers differ or are wrong.
Overall:
The reusable payload to LEO has been cut by about half, from 300t to 150t. The payload to Mars has gone from 450t to 150t, and the goal for the number of people per flight has gone from 100+ to 100.
Raptor
The Raptor engine is further along in development than last time, and has been slightly downscaled to match the smaller BFR. This is because it is harder to make engines throttle to lower percentages of full thrust, which they would need to do to land safely, and still have redundancy.
"The engine thrust dropped roughly in proportion to the vehicle mass reduction from the first IAC talk" (source)
Elon Musk said in the presentation that the Isp had the potential to be increased by 5-10 seconds, and the chamber pressure by 50 bar, which would make the 2017 Raptor have the same stats as the 2016 one.
(2017)(2016)
250bar vs 300bar
Deep throttle to 20%/ 20%
Vac
Exit diameter 2.4m/ expansion ratio 200
thrust 1,900kn/ 3,500kn
isp 375s/ 382s
SL
exit dia 1.3m/ expansion ratio 40
Thrust (SL) 1,700kn/ (SL) 3,050kn
Isp (SL) 330s, (Vac) 356s/ Isp (SL) 334s, (Vac) ???
BFB
Dimensions:
(2017)(2016)
58m x 9m/ 77.5m x 12m
Mass is tricky, as the video of the presentation gives exact values, while the PDF of the slides give a different exact value for thrust, and no numbers for mass. For this post, I'll use the numbers described in the presentation. Dry and prop masses are based on ratios from the 2016 BFB.
Mass (dry, prop, wet, prop mass fraction[prop/wet])
275t, 6700t, 6975t, ~0.96
to
~126.75t, ~3,088.25t, 3215t, ~0.96
Engines:
(2017)(2016)
31/ 42
(Apparently the important scientific and fictional reasons weren't that important, however the PDF does not say how many engines it has, which means that it is likely changing.)
Thrust:
This is weird. The talk showed a thrust of 5400 tons, which is equal to 48040.79kn, which is how much thrust 28 and a quarter SL raptors produce ASL. It's possible that this is a mistake, as there was also a mistake in the interior volume of the BFS, but it's also possible that it was from a different version of raptor.
In the PDF, this was changed to 52,700kn, which is exactly what 31 SL raptors produce ASL, lending weight to the theory that this was a typo. I'm going to assume that it was a typo.
(2017)(2016)
52,700kn/ 128mn (SL) 138mn (Vac)
Delta-V
Vac:
11288.2m/s / ???
SL:
10463.8m/s / 10590.47m/s
(note: this seems odd, as the thrust and fuel mass are roughly double in the 2016 BFR, but remember that Delta-V is a function of propellant ratio and Isp, and I assume the propellant ratio to be the same. Basically, scaling up or down a spaceship design just changes how much extra payload affects Delta-V)
BFS:
The spaceship's largest change, aside from size, is in shape. The 2016 IAC spaceship had a much more complex body shape, with 3 fins that blended into the overall shape of the ship. The 2017 ship is much simpler, with a simple cylindrical shape and a single set of delta wings. This was done to avoid building a "box in a box" (source)
The BFS now refuels via a connection at the end of the ship, rather than at the side.
Size
(2017)(2016)
48m x 9m/ 49.5m x 17m (max) 12m (base)
Mass
Using the masses from the interplanetary ship in the 2016 numbers.
(2017)(2016)(dry, prop, wet, prop mass fraction[prop/wet])
85t, 1,100t, 1,185, ~.93 / 150t, 1,950t, 2100t, ~0.93
(In the presentation, Elon Musk said that the current design has the dry mass as 75t, but mass growth would likely occur)
Delta-V
(All vac)
9689.63m/s / 9886.28 m/s
Engines
4 Vac, 2 SL/ 6 Vac, 3 SL
Thrust
Vac
7,600kn/ 21,000kn
SL
3,400kn/ 9,150kn
However, Elon Musk said here that a 3rd medium area Raptor was added to the BFS since the IAC. I don't know what form this would take, or how this would fit on the BFS.
Monday, October 30, 2017
Sunday, July 30, 2017
Interesting space videos
Today I'm taking a break from the alternate ITS missions series, as I'm not yet finished with it. So I've going to post a couple of videos about spaceflight I've found.
First, a video from Scott Manley, about reused spacecraft.
This video is a timelapse of the night sky, but with the motion of the Earth corrected, showing how the Earth moves relative to the stars. Bonus video.
This shows six daylight Falcon 9 landings, timed so they all land at the same time.
And this, from Tom Scott, is a video on how spaceflight hardware and zero-g experiments that do not need enough time in zero-g to require a parabolic flight are tested.
First, a video from Scott Manley, about reused spacecraft.
This video is a timelapse of the night sky, but with the motion of the Earth corrected, showing how the Earth moves relative to the stars. Bonus video.
This shows six daylight Falcon 9 landings, timed so they all land at the same time.
And this, from Tom Scott, is a video on how spaceflight hardware and zero-g experiments that do not need enough time in zero-g to require a parabolic flight are tested.
Sunday, July 2, 2017
Alternate ITS missions part 3
This post got too long again, so I'm splitting it into yet another section. This will cover how to calculate interplanetary trajectories.
All orbits are conic sections. circles, ellipses, parabolas, or hyperbolas. Circles and ellipses are the orbits that a spacecraft is in when it is in a stable orbit around a planet. Parabolas are when the spacecraft is going at exactly escape velocity, and hyperbolas are when the spacecraft is going faster than escape velocity.
For ellipses (assuming circles are a type of ellipse) the planet the spacecraft is orbiting (or the primary) is always at one of the two foci. The major and minor axes are the longest and the shortest lines that can be drawn through the ellipse, respectively. The semi-major or minor axes are half of the major or minor axes.
The periapsis is the lowest point on an orbit, and the apoapsis is the highest. They are always 180 degrees apart. The closer to the periapsis, the faster the satellite travels, and the closer to apoapsis, the slower.
A quick aside on nodes: nodes are the points at which the orbit passes the plane of the equator or ecliptic. The ascending node is the one the spacecraft passes travelling from south to north, and the descending node is the one the spacecraft passes travelling from north to south.
Eccentricity is the distance between the foci divided by the length of the major axis (always between 0 and 1 except for hyperbolas), and is essentially a measure of how "squished" the ellipse is. 0 is a circle, 1 is parabola, > 1 is a hyperbola.
Inclination is how tilted the orbit is from the equatorial plane (or the ecliptic) of the primary. An inclination of less than 90 degrees indicates a orbit in the same direction as the spin of the primary, or prograde, an inclination of greater than 90 degrees means that the orbit is in the opposite direction of the spin, or retrograde.
I'm ignoring how to calculate these, as that would make this post much longer.
So, let's imagine you want to go from one circular orbit to another. What is the most efficient way of doing this? The answer is a Hohmann transfer.
A Hohmann transfer is the most efficient way to go from a coplanar orbit to another. The image shows a transfer with two circular orbits, however it is the most efficient for any two coplanar orbits where one cannot be reached from the other with only a single burn. For example, in the image, 1 and 3 require a hohmann transfer, while 1 and 2 or 2 and 3 only need a single burn.
The Delta-V required for a hohmann transfer is the sum of the Delta-V required for the burn to go from 1 to 2, and the Delta-V required to go from 2 to 3. It is the same in the opposite direction (3 to 1 = 3 to 2 + 2 to 1), and the burns are the same, except with the spacecraft burning its engines in the opposite direction.
The transfer orbit between two planets is the same thing, with the transfer orbit going from one planet to another, except for the fact that you start in orbit around one of those planets.
Transfer windows are times when the orbits of the planets align such that a spacecraft on a transfer orbit reaches apoapsis (or periapsis) at the same time as the destination planet moves through the part of the orbit that intersects the transfer orbit.
It isn't very important to calculate transfer windows for this post, as they do occur regularly for most planets.
Delta-V for an interplanetary transfer is less than you might expect, as the Oberth effect comes into play on the escape burn. The Oberth effect is, in short, an effect that makes Delta-V count for more when near a large mass. It is why the Delta-V required for a hohmann transfer to GEO is more than the Delta-V required for escape velocity.
The Oberth effect is the effect of the extra velocity of a circular orbit around earth has compared to a circular orbit around the sun at earth's altitude. As the satellite orbits around the earth, it traces out a spiral pattern relative to the sun. When it is orbiting in the same direction around the earth that the earth is orbiting around the sun, the velocity relative to the sun is increased. It isn't as simple as adding the velocities, but we'll get into that later.
On arrival at the destination planet, the Delta-V can be further reduced with aerobraking or aerocapture. This is where the you use the atmosphere of the planet you are arriving at to slow down or speed up without using any fuel. This only works on planets with atmospheres, though, and you cannot capture into a circular orbit without a burn, although it is much less with most of the velocity bled off already.
A very useful equation is the Vis-visa equation:
v^2 = GM(2/r - 1/a)
Where v is the speed of the spacecraft at any point in a orbit, G is the gravitational constant, M is the mass of the primary, r is the distance between the spacecraft and the primary at the point at which you wish to know velocity, and a is the semi-major axis.
The calculations for hohmann transfer orbits assume instantaneous impulses, which lessens delta-v requirements. To compensate, I will multiply the estimate by 1.1.
To calculate Delta-V requirements, you simply need to get the difference in velocities between the orbits at the correct points. This is why the Vis-visa equation is so useful.
We apply this to a hohmann transfer from one planet to another, getting the velocities of the transfer at apoapsis and periapsis. Then, we calculate the velocity of the starting orbit, and calculate the required Delta-V to go to the transfer orbit, accounting for the Oberth effect.
Reversing the rocket equation will then let us calculate payload.
Earth to Venus example
Earth has a mass of 5.9723 * 10^24 kg. It has a semi-major axis of 149.6 * 10^9 meters. Its orbital radius goes from 147.09 * 10^9 m to 152.1 * 10^9 m.
Venus has a mass of 4.8675 * 10^24 kg. It has a semi-major axis of 108.21 * 10^9 m. Its orbital radius goes from 107.48 * 10^9 m to 108.94 * 10^9 m.
The sun has a mass of 1988500 * 10^24 kg.
A transfer orbit between them depends on the position of the planets at departure and arrival. For this example, the position of Venus does not matter, as aerobraking will work at either of those speeds.
Earth should be at apoapsis, because in this case we want to lose, rather than gain, velocity as Venus is in a lower orbit than Earth.
Let's begin. First, the velocity of Earth around the Sun at apoapsis. We don't really need to calculate this, as it is well known: 29290 m/s.
The transfer orbit is an elliptical orbit with its apoapsis at Earth, and its periapsis at Venus. This gives a semi-major axis of:
(Radius of Earth's orbit at departure + Radius of Venus' orbit at arrival)/2.
Or (152.1 * 10^9 m + 108.94 * 10^9 m)/2 = 130.52 * 10^9 m.
Velocity at apoapsis:
v^2 = 1.327 * 10^20 (2/152.1 * 10^9 - 1/130.52 * 10^9)
v = 26985 m/s
At periapsis:
v^2 = 1.327 * 10^20 (2/108.94 * 10^9 - 1/130.52 * 10^9)
v = 37676 m/s
The velocity of the spaceship in Earth orbit can be calculated from the altitude, and the altitude is as low as possible to take maximum advantage of the Oberth effect. I'm going to guess around 200 km.
v^2 = 3.986 * 10^14 (2/6578000 - 1/6578000)
v = 7784 m/s.
To figure out how much of an advantage the Oberth effect gives us, we use something called hyperbolic excess velocity. This is how much extra velocity the spaceship will have when it reaches infinite distance on a hyperbolic escape orbit. If you reach exactly escape velocity, HEV will be zero. The desired HEV is equal to the orbital velocity of the primary - the velocity of the transfer orbit at the point where it begins so when the spacecraft escapes the primary's gravity, the remaining velocity puts it on the transfer orbit.
The equation for HEV is HEV^2 = v^2 - ve^2
Our desired HEV is 29290 - 26985 = 2305.
Rearranging the equation:
v^2 =2305^2 + 11008^2.
(as escape velocity is equal to the velocity of a circular orbit at that altitude times √2)
So v = 11246 m/s.
11246 - 7784 = 3462.
Here you can see the clear advantage of this over a escape burn (3224) + a burn outside of Earth's sphere of influence to put the spacecraft on a transfer orbit (2305), which would be 2067 m/s more.
Venus is inclined 3.39 degrees off of Earth, however this amounts to at most a few m/s if it is corrected for at the very start of the transfer. For this kind of loose approximation, we can ignore this.
So, multiplying by 1.1, we have a Delta-V of 3808 m/s.
Next post, we will calculate transfers for other missions, including approximate landing Delta-V, and use the reverse rocket equation to figure out how much payload can be carried to them.
All orbits are conic sections. circles, ellipses, parabolas, or hyperbolas. Circles and ellipses are the orbits that a spacecraft is in when it is in a stable orbit around a planet. Parabolas are when the spacecraft is going at exactly escape velocity, and hyperbolas are when the spacecraft is going faster than escape velocity.
For ellipses (assuming circles are a type of ellipse) the planet the spacecraft is orbiting (or the primary) is always at one of the two foci. The major and minor axes are the longest and the shortest lines that can be drawn through the ellipse, respectively. The semi-major or minor axes are half of the major or minor axes.
The periapsis is the lowest point on an orbit, and the apoapsis is the highest. They are always 180 degrees apart. The closer to the periapsis, the faster the satellite travels, and the closer to apoapsis, the slower.
A quick aside on nodes: nodes are the points at which the orbit passes the plane of the equator or ecliptic. The ascending node is the one the spacecraft passes travelling from south to north, and the descending node is the one the spacecraft passes travelling from north to south.
Eccentricity is the distance between the foci divided by the length of the major axis (always between 0 and 1 except for hyperbolas), and is essentially a measure of how "squished" the ellipse is. 0 is a circle, 1 is parabola, > 1 is a hyperbola.
Inclination is how tilted the orbit is from the equatorial plane (or the ecliptic) of the primary. An inclination of less than 90 degrees indicates a orbit in the same direction as the spin of the primary, or prograde, an inclination of greater than 90 degrees means that the orbit is in the opposite direction of the spin, or retrograde.
I'm ignoring how to calculate these, as that would make this post much longer.
So, let's imagine you want to go from one circular orbit to another. What is the most efficient way of doing this? The answer is a Hohmann transfer.
Image from Leafnode
The Delta-V required for a hohmann transfer is the sum of the Delta-V required for the burn to go from 1 to 2, and the Delta-V required to go from 2 to 3. It is the same in the opposite direction (3 to 1 = 3 to 2 + 2 to 1), and the burns are the same, except with the spacecraft burning its engines in the opposite direction.
The transfer orbit between two planets is the same thing, with the transfer orbit going from one planet to another, except for the fact that you start in orbit around one of those planets.
Transfer windows are times when the orbits of the planets align such that a spacecraft on a transfer orbit reaches apoapsis (or periapsis) at the same time as the destination planet moves through the part of the orbit that intersects the transfer orbit.
It isn't very important to calculate transfer windows for this post, as they do occur regularly for most planets.
Delta-V for an interplanetary transfer is less than you might expect, as the Oberth effect comes into play on the escape burn. The Oberth effect is, in short, an effect that makes Delta-V count for more when near a large mass. It is why the Delta-V required for a hohmann transfer to GEO is more than the Delta-V required for escape velocity.
The Oberth effect is the effect of the extra velocity of a circular orbit around earth has compared to a circular orbit around the sun at earth's altitude. As the satellite orbits around the earth, it traces out a spiral pattern relative to the sun. When it is orbiting in the same direction around the earth that the earth is orbiting around the sun, the velocity relative to the sun is increased. It isn't as simple as adding the velocities, but we'll get into that later.
On arrival at the destination planet, the Delta-V can be further reduced with aerobraking or aerocapture. This is where the you use the atmosphere of the planet you are arriving at to slow down or speed up without using any fuel. This only works on planets with atmospheres, though, and you cannot capture into a circular orbit without a burn, although it is much less with most of the velocity bled off already.
A very useful equation is the Vis-visa equation:
v^2 = GM(2/r - 1/a)
Where v is the speed of the spacecraft at any point in a orbit, G is the gravitational constant, M is the mass of the primary, r is the distance between the spacecraft and the primary at the point at which you wish to know velocity, and a is the semi-major axis.
The calculations for hohmann transfer orbits assume instantaneous impulses, which lessens delta-v requirements. To compensate, I will multiply the estimate by 1.1.
To calculate Delta-V requirements, you simply need to get the difference in velocities between the orbits at the correct points. This is why the Vis-visa equation is so useful.
We apply this to a hohmann transfer from one planet to another, getting the velocities of the transfer at apoapsis and periapsis. Then, we calculate the velocity of the starting orbit, and calculate the required Delta-V to go to the transfer orbit, accounting for the Oberth effect.
Reversing the rocket equation will then let us calculate payload.
Earth to Venus example
Earth has a mass of 5.9723 * 10^24 kg. It has a semi-major axis of 149.6 * 10^9 meters. Its orbital radius goes from 147.09 * 10^9 m to 152.1 * 10^9 m.
Venus has a mass of 4.8675 * 10^24 kg. It has a semi-major axis of 108.21 * 10^9 m. Its orbital radius goes from 107.48 * 10^9 m to 108.94 * 10^9 m.
The sun has a mass of 1988500 * 10^24 kg.
A transfer orbit between them depends on the position of the planets at departure and arrival. For this example, the position of Venus does not matter, as aerobraking will work at either of those speeds.
Earth should be at apoapsis, because in this case we want to lose, rather than gain, velocity as Venus is in a lower orbit than Earth.
Let's begin. First, the velocity of Earth around the Sun at apoapsis. We don't really need to calculate this, as it is well known: 29290 m/s.
The transfer orbit is an elliptical orbit with its apoapsis at Earth, and its periapsis at Venus. This gives a semi-major axis of:
(Radius of Earth's orbit at departure + Radius of Venus' orbit at arrival)/2.
Or (152.1 * 10^9 m + 108.94 * 10^9 m)/2 = 130.52 * 10^9 m.
Velocity at apoapsis:
v^2 = 1.327 * 10^20 (2/152.1 * 10^9 - 1/130.52 * 10^9)
v = 26985 m/s
At periapsis:
v^2 = 1.327 * 10^20 (2/108.94 * 10^9 - 1/130.52 * 10^9)
v = 37676 m/s
The velocity of the spaceship in Earth orbit can be calculated from the altitude, and the altitude is as low as possible to take maximum advantage of the Oberth effect. I'm going to guess around 200 km.
v^2 = 3.986 * 10^14 (2/6578000 - 1/6578000)
v = 7784 m/s.
To figure out how much of an advantage the Oberth effect gives us, we use something called hyperbolic excess velocity. This is how much extra velocity the spaceship will have when it reaches infinite distance on a hyperbolic escape orbit. If you reach exactly escape velocity, HEV will be zero. The desired HEV is equal to the orbital velocity of the primary - the velocity of the transfer orbit at the point where it begins so when the spacecraft escapes the primary's gravity, the remaining velocity puts it on the transfer orbit.
The equation for HEV is HEV^2 = v^2 - ve^2
Our desired HEV is 29290 - 26985 = 2305.
Rearranging the equation:
v^2 =2305^2 + 11008^2.
(as escape velocity is equal to the velocity of a circular orbit at that altitude times √2)
So v = 11246 m/s.
11246 - 7784 = 3462.
Here you can see the clear advantage of this over a escape burn (3224) + a burn outside of Earth's sphere of influence to put the spacecraft on a transfer orbit (2305), which would be 2067 m/s more.
Venus is inclined 3.39 degrees off of Earth, however this amounts to at most a few m/s if it is corrected for at the very start of the transfer. For this kind of loose approximation, we can ignore this.
So, multiplying by 1.1, we have a Delta-V of 3808 m/s.
Next post, we will calculate transfers for other missions, including approximate landing Delta-V, and use the reverse rocket equation to figure out how much payload can be carried to them.
Wednesday, May 31, 2017
Alternate ITS missions part 2
Unfortunately, this post will have to be broken into three parts, since I started looking into how life support worked, and the post got really long. This part will mostly be about life support, and the next one will have details on actual missions.
There are two primary ways to support humans in space with oxygen, food, and water: ecological, and chemical. Ecological uses plants to reprocess human waste, carbon dioxide, and trace nutrients into food, and humans to process food and oxygen into what the plants need, making a closed loop, as long as you have sunlight.
The largest problem with this is that the amount of plants required to support one person's oxygen needs would produce about half the required food for that person. If you have enough plants to provide enough food, half of them die off due to lack of carbon dioxide.
Chemical life support uses chemical reactors to crack carbon dioxide into oxygen, and extract usable water from sweat and urine. The problem with this is that you must pack all the food you will need.
Storage of oxygen is simple, as you can use boiloff from the main oxidizer tanks.
The air on Earth at sea level is at 101kpa. The atmosphere is 21% oxygen, so the partial pressure of oxygen is 21.21kpa.
Partial pressure, in a mix of gases, is how much pressure one of the gases would be under if all other gases were removed.
The minimum safe partial pressure of oxygen is 16kpa, and the maximum is somewhere between 50 and 100 kpa for short periods of time (a few hours). For long term breathing atmosphere, you want to stay under 50 kpa, ideally at about 21 kpa.
A high pressure pure oxygen environment is very dangerous because of fire, the Apollo 1 fire happened in a pure oxygen environment at 115kpa, this is what wood burning in a 50 kpa partial pressure environment looks like: https://youtu.be/_JkHB1hV7Hw?t=1m After that all US manned spacecraft used a 79% nitrogen/21% oxygen environment at about 101kpa.
So why was the Apollo 1 command module pressurized to 115 kpa? Because it was designed to hold pressure in, not out. During launch, the pressure would have gradually been reduced to 34 kpa. I don't know whether the ITS will use a nitrogen/oxygen mix, or pure oxygen which is always kept at safe pressures, but either way it will take about 0.85kg/day of oxygen, as nitrogen is not consumed when inhaled.
A larger problem than oxygen is carbon dioxide. Every astronaut exhales about 1 kg of it a day, and if it rises above 0.5% concentration it can become a serious danger. However, 0.273 kg of that is carbon, so most of it can be breathed again if it is cracked back into carbon and oxygen.
There are two options for dealing with carbon dioxide, cracking and scrubbing. Scrubbing uses a catalyst, which removes carbon dioxide from the air, producing water. It can be cleaned and used again by blowing hot air through it for 10 hours, however, you lose all that oxygen locked in the carbon dioxide.
A chemical reactor would use the sabatier reaction, like the ISS.
CO2 + 4H2 → CH4 + 2H2O
Exhaled CO2 would be processed with hydrogen to produce CH4 (methane) and water. The water can be electrolyzed into H2 and O2, and the CH4 can be pyrolized into C and H2. The hydrogen outputs can be fed back into the sabatier reactor, closing the loop. It would probably not be perfectly efficient, but since you only need about 180 grams of H2 for every kg or CO2 processed, even at 90% efficiency you would only need to add 18 grams of hydrogen a day.
However, every crew member exhales 1 kg of carbon dioxide/day, but only 0.727 kg of that is oxygen. That means that you have to add 108 grams of oxygen every day.
Astronauts need to drink about 3.9 kg of water every day, some mixed in with their food, if it isn't freeze-dried, and about 26 kg/day of water for personal hygiene. Most of the waste water (grey water, human waste, sweat) can be distilled and filtered. I estimate that about 0.1 kg of water would be lost in recovery every day.
Astronauts on the ISS eat about 2.5 kg of food every day, which consists of a mix of different kinds of food, mostly freeze-dried or otherwise stabilized. Some of it is fresh.
For the purposes of this, we can assume that all non-reusable waste (packaging, filters, solid human waste) is non-existent, as it can be dumped overboard between burns. If you wished to prevent the build up of trash in solar orbit you could dump it only while on a collision course with a planet, however that would likely only be a problem on trips that would be taken frequently.
Weights required:
5 tons for storage, processing equipment, etc.100 grams of water x mission length in days18 grams of H2 x mission length in days108 grams of oxygen x mission length in days x crew members2.5 kg of food x mission length in days x crew members
In a CELSS (Closed (or Controlled) ecological life support system), rather than using chemical reactors, the life support loop is fully closed, like Earth's ecosystem.
If you could get a CELSS working properly, it has the potential to be much more mass efficient than chemical life support system, with the crossover point for efficiency at around 2 years, according to Rocketpunk Manifesto. However, in their current state, they have some problems. For many plants, the amount of plants needed to provide food for astronauts is different for the amount need to provide oxygen, so the imbalance would cause an excess of oxygen, which would kill some of the plants, and then there wouldn't be enough food.
One of the most promising crops is Spirulina, a type of blue-green algae that, Marshall T. Savage claims in The Millennial Project, could close the life support loop almost fully with only 6 liters of algae per person (about 6.6 kg). (Note: I haven't read that book, this is taken from Atomic Rockets)
Even if you needed twice that much per person it would be very impressive.
For the purposes of this analysis, however, I will stick to existing methods of food production, as the goal is to use the fact that the ITS would be flying regularly to study a mission that could be much cheaper than a spacecraft designed only for the purpose of that mission.
In the next part we will finally look at the actual mission profiles.
There are two primary ways to support humans in space with oxygen, food, and water: ecological, and chemical. Ecological uses plants to reprocess human waste, carbon dioxide, and trace nutrients into food, and humans to process food and oxygen into what the plants need, making a closed loop, as long as you have sunlight.
The largest problem with this is that the amount of plants required to support one person's oxygen needs would produce about half the required food for that person. If you have enough plants to provide enough food, half of them die off due to lack of carbon dioxide.
Chemical life support uses chemical reactors to crack carbon dioxide into oxygen, and extract usable water from sweat and urine. The problem with this is that you must pack all the food you will need.
Storage of oxygen is simple, as you can use boiloff from the main oxidizer tanks.
The air on Earth at sea level is at 101kpa. The atmosphere is 21% oxygen, so the partial pressure of oxygen is 21.21kpa.
Partial pressure, in a mix of gases, is how much pressure one of the gases would be under if all other gases were removed.
The minimum safe partial pressure of oxygen is 16kpa, and the maximum is somewhere between 50 and 100 kpa for short periods of time (a few hours). For long term breathing atmosphere, you want to stay under 50 kpa, ideally at about 21 kpa.
A high pressure pure oxygen environment is very dangerous because of fire, the Apollo 1 fire happened in a pure oxygen environment at 115kpa, this is what wood burning in a 50 kpa partial pressure environment looks like: https://youtu.be/_JkHB1hV7Hw?t=1m After that all US manned spacecraft used a 79% nitrogen/21% oxygen environment at about 101kpa.
So why was the Apollo 1 command module pressurized to 115 kpa? Because it was designed to hold pressure in, not out. During launch, the pressure would have gradually been reduced to 34 kpa. I don't know whether the ITS will use a nitrogen/oxygen mix, or pure oxygen which is always kept at safe pressures, but either way it will take about 0.85kg/day of oxygen, as nitrogen is not consumed when inhaled.
A larger problem than oxygen is carbon dioxide. Every astronaut exhales about 1 kg of it a day, and if it rises above 0.5% concentration it can become a serious danger. However, 0.273 kg of that is carbon, so most of it can be breathed again if it is cracked back into carbon and oxygen.
There are two options for dealing with carbon dioxide, cracking and scrubbing. Scrubbing uses a catalyst, which removes carbon dioxide from the air, producing water. It can be cleaned and used again by blowing hot air through it for 10 hours, however, you lose all that oxygen locked in the carbon dioxide.
A chemical reactor would use the sabatier reaction, like the ISS.
CO2 + 4H2 → CH4 + 2H2O
Exhaled CO2 would be processed with hydrogen to produce CH4 (methane) and water. The water can be electrolyzed into H2 and O2, and the CH4 can be pyrolized into C and H2. The hydrogen outputs can be fed back into the sabatier reactor, closing the loop. It would probably not be perfectly efficient, but since you only need about 180 grams of H2 for every kg or CO2 processed, even at 90% efficiency you would only need to add 18 grams of hydrogen a day.
However, every crew member exhales 1 kg of carbon dioxide/day, but only 0.727 kg of that is oxygen. That means that you have to add 108 grams of oxygen every day.
Astronauts need to drink about 3.9 kg of water every day, some mixed in with their food, if it isn't freeze-dried, and about 26 kg/day of water for personal hygiene. Most of the waste water (grey water, human waste, sweat) can be distilled and filtered. I estimate that about 0.1 kg of water would be lost in recovery every day.
Astronauts on the ISS eat about 2.5 kg of food every day, which consists of a mix of different kinds of food, mostly freeze-dried or otherwise stabilized. Some of it is fresh.
For the purposes of this, we can assume that all non-reusable waste (packaging, filters, solid human waste) is non-existent, as it can be dumped overboard between burns. If you wished to prevent the build up of trash in solar orbit you could dump it only while on a collision course with a planet, however that would likely only be a problem on trips that would be taken frequently.
Weights required:
5 tons for storage, processing equipment, etc.100 grams of water x mission length in days18 grams of H2 x mission length in days108 grams of oxygen x mission length in days x crew members2.5 kg of food x mission length in days x crew members
In a CELSS (Closed (or Controlled) ecological life support system), rather than using chemical reactors, the life support loop is fully closed, like Earth's ecosystem.
If you could get a CELSS working properly, it has the potential to be much more mass efficient than chemical life support system, with the crossover point for efficiency at around 2 years, according to Rocketpunk Manifesto. However, in their current state, they have some problems. For many plants, the amount of plants needed to provide food for astronauts is different for the amount need to provide oxygen, so the imbalance would cause an excess of oxygen, which would kill some of the plants, and then there wouldn't be enough food.
One of the most promising crops is Spirulina, a type of blue-green algae that, Marshall T. Savage claims in The Millennial Project, could close the life support loop almost fully with only 6 liters of algae per person (about 6.6 kg). (Note: I haven't read that book, this is taken from Atomic Rockets)
Even if you needed twice that much per person it would be very impressive.
For the purposes of this analysis, however, I will stick to existing methods of food production, as the goal is to use the fact that the ITS would be flying regularly to study a mission that could be much cheaper than a spacecraft designed only for the purpose of that mission.
In the next part we will finally look at the actual mission profiles.
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