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Exoscientist

Propellant depot based Mars architecture.

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Key points about a propellant depot based Mars architecture, once the propellant depots are in place at both departure and arrival points:

1.)A single medium-lift booster first stage, Falcon 9, Atlas V, Delta IV, etc., delivered empty to orbit can then do ALL the propulsion from LEO departure, to Mars orbit insertion, to Mars landing, to Mars liftoff, to return to Earth.

No Saturn V, Constellation, Ares V, SLS, Mars Colonial Transport, or even Falcon Heavy required. The required boosters are already existing IF those propellant depots are already in place.

2.)SpaceX has shown that you can do reentry burns in the hypersonic airstream with the F9 first stage reuse tests. Then the problem of landing large masses on Mars is solved by doing a fully propulsive burn to Mars landing once that one, single stage is refueled in Mars orbit.

3.)That one single mid-lift stage could also be used to make an approx. 30 day flight to Mars. No VASIMR, solar electric, or nuclear propulsion required. However, very high reentry velocity heat shields, ca. 20 km/s instead of ca. 6 km/s, would need to be developed for this.

4.)The most important point of all: getting the propellant depots to cislunar orbit is easy using near Earth asteroids. You don't need to use the Moon's proposed water ice deposits or develop a manned lunar base. This was the most surprising calculation of all: a single Centaur upper stage, of ca. 20 metric ton(mT) gross mass, could drag a 500 mT asteroid to cislunar space.

See:

Propellant depots for interplanetary flight.

http://exoscientist.blogspot.com/2015/08/propellant-depots-for-interplanetary.html

Bob Clark

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Until someone comes up with a proven method for mining NEOs (which is not likely for the next 10 or 20 years at least), propellant depots still require that the propellant itself is sent from the surface. So scratch that if we are talking about feasible projects instead of science fiction.

If a 100-ton MTV needs 500 tons of propellant to go to Mars, then you still need to launch the 100-ton MTV empty plus its 500 tons of propellant. But if the MTV has its own tanks, then why bother with a depot? Just use your MTV as a depot while waiting for the next window to Mars.

I'd take it even further: why even put tanks on your MTV in the first place? The fuel will be coming up in disposable tanker vehicles anyway, which will basically just be a tank with a docking ring and RCS. So scrap the whole depot idea, and just dock a bunch of tankers to the MTV. All the MTV needs is a hab and engines, and you can dump the tanks as you go and plug in new ones before each journey.

The only case where a depot might increase efficiency is if you have regular flights between Mars and the Earth and you need a fast turnaround, but that isn't happening anytime soon either...

Edited by Nibb31

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If the MTV uses cryogens and the mission architecture assumes no martian ISRU, then it seems likely that propellant depots will developmentally precede the MTV. Zero boiloff, cryogenic fluid transfer, acquisition and mass gauging technologies are still at low technology readiness levels. A depot could double as a technology demonstrator mission, and enable the accumulation of a stockpile of propellant in orbit in parallel to launches of the MTV (rather than waiting till it is fully assembled, and checked out, on orbit). This parallelization should help reduce mission schedule risk (likely to be substantial for such a complex mission anyway). On the other hand using depots adds an additional logistical cycle to the process; this means more chance for something to break or for leaks to occur. In the near term propellant depots have commercial applications, and are frankly much more likely to exist in the near term than a Mars mission.

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Zero boiloff

Why zero boiloff?

Boiling point -200..250° C, equlibrium temperature for Venus..Mars orbits -100..+50°C.

Would boil off and how!

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Why zero boiloff?

Boiling point -200..250° C, equlibrium temperature for Venus..Mars orbits -100..+50°C.

Would boil off and how!

Because it's cryogenic hydrogen with a saturation temperature of 20K! That is really cold! If you don't refrigerate that stuff or at least insulate it realllyy well it'll all boiling away and your tanks will be empty long before you reach Mars orbit, leaving you stranded (or worse, you can't even brake into orbit without aerocapture).

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...

4.)The most important point of all: getting the propellant depots to cislunar orbit is easy using near Earth asteroids. You don't need to use the Moon's proposed water ice deposits or develop a manned lunar base. This was the most surprising calculation of all: a single Centaur upper stage, of ca. 20 metric ton(mT) gross mass, could drag a 500 mT asteroid to cislunar space.

The ideal asteroid candidates for retrieval have low delta-v requirements to move to cislunar space, such as to L2 or lunar capture. For one known asteroid it's particularly low, 2008Hu4, at only 170 m/s. It's orbital parameters are given here:

2008 HU4

2nb7ng0.jpg

http://ssd.jpl.nasa.gov/sbdb.cgi?sstr=2008%20HU4;orb=1

The closest approach is in April, 2016. How would a chemical propulsion transfer look that provided the needed 170 m/s? For the chemical propulsion transfer you may assume the delta-v is provided in a single short impulse. How long would it take for the transfer at the closest approach?

Note that when you click on the "Close-Approach Data" link on that page for 2008HU4 it gives the V-infinity with respect to the Earth as 1.28 km/s. This means their relative speed before it is effected by Earth's gravity and speeded up. This is relevant because remember we don't want to put it in Earth orbit but put it in lunar orbit or at L2:

Asteroid Redirect Mission.

https://en.wikipedia.org/wiki/Asteroid_Redirect_Mission

Then since the Moon's orbital speed around Earth is 1.1 km/s, it could be a small relative speed between the asteroid and the Moon. But this would depend on the position of the Moon when the asteroid makes its closest approach.

Most discussions of ARM just look at solar electric propulsion(SEP) because it would give a smaller mission size. But the thing is when the delta-v is so small as 170 m/s you could move a 500 metric ton (mT) asteroid with just a single Centaur upper stage, at ca. 20 mT gross mass.

That's the case I want to look at because of the 5 to 10 year transfer time for the SEP case. It should be a shorter transfer time when using chemical propulsion. If you want, you can just calculate what would be the transfer time for the asteroid to get within the Moon's distance of the Earth. We can assume we use ballistic capture or small delta-v burn when it comes close to the Moon to put it in lunar orbit or at L2.

Bob Clark

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Tater on the "39 days to Mars possible now with nuclear-powered VASIMR" thread noted that delta-v from Phobos or Deimos is lower than from the Moon and therefore it would be easier to get propellant from there. It is indeed the case that at least for Phobos its low density is believed to be due to large amounts of water/ice content.

Looking at this chart of delta-v's to the Mars system I was surprised how low was the delta-v to get from Phobos or Deimos to Earth assuming you use full aerobraking on arrival at Earth:

Mars-Moon-Earth+Delta-v.png

You see to lift-off from Deimos, exit out of Deimos orbit around Mars, and be put on a transfer trajectory towards Earth would require delta-v's of .7 + .2 +.9 km/s = 1.8 km/s. And for Phobos, it would be .5 + .3 + .2 + .9 km/s = 1.9 km/s. After this you assume you do aerobraking on return at Earth. This is less than that for the Moon.

Also quite interesting is if you add up the required delta-v's to get to the moons of Mars and then return, the total is also surprisingly low, 7.4 km/s for Deimos and 7.5 km/s for Phobos. This means it would be quite easy in delta-v terms to do a sample return mission from them. It could be launched by the Falcon 9 using existing upper stages to serve as the in-space stages. On an up coming blog post I'll describe this.

About the return flight though, when delta-v's to and from Mars are quoted it's usually for a Hohmann transfer trajectory for when Mars and Earth are at closest approach. So when the sample return mission was to head back to Earth it would have to wait two years to have the low delta-v requirements of the close approach to recur. But perhaps someone on Kerbal could do the calculation for the required delta-v if you departed back to Earth soon after arrival. It would be larger, but likely still doable by a Falcon 9 launch.

Bob Clark

Edited by Exoscientist

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Why zero boiloff?

Boiling point -200..250° C, equlibrium temperature for Venus..Mars orbits -100..+50°C.

Would boil off and how!

The equilibrium temperature depends on the radiative properties of the spacecraft. Your temperatures are reasonable for planet surfaces, but for an artificial material with very low solar-emissivity, Teq can be much colder than -100 °C. Moreso, there's the trick that you can stack several layers of radiating surfaces, to get a combined, effective emissivity much lower than any physical material's. For a famous example, the James Webb telescope has a sunshield with five layers, εeff = 0.00038 (!!), and Teq = 40 K (-233 °C) on the cold side. That temperature is achieved solely by passive, radiative cooling.

3CpTttx.jpgQMbeviY.jpg

http://jwst.nasa.gov/sunshield.html

In theory, you can get Teq so cold, that it's below the boiling point of liquid hydrogen (20 K at 1 atm). Alternatively, it can be slightly warmer, and the tank can be kept artificially cold by active refrigeration. Either way, no hydrogen is vented into space: it's "zero boil-off" (ZBO).

Thermal analysis results show that it is possible to store LOX/LH2 at reasonable tank pressures using only passive radiation cooling when the field of view of propellant tanks can be kept clear of warm planetary bodies. This situation is typical of interplanetary cruise, spacecraft orbiting bodies with low effective blackbody temperatures, and spacecraft with very short stay times near the target planet, asteroid, or comet. Passive storage was accomplished using a combination of sun shades, spacecraft configuration considerations, spacecraft pointing constraints, and the low conductance Passive Orbit Displacement Strut (PODS). Actively cooled designs use cryocoolers and mechanically pumped fluid loops to reject heat from the propellant tanks.

4va2TAH.png

http://develop.nttc.edu/sbipp/technologyportfolios/portfolios/ISS-Propellant_T-S/Archive%5C20060042833.pdf

- - - Updated - - -

Most discussions of ARM just look at solar electric propulsion(SEP) because it would give a smaller mission size. But the thing is when the delta-v is so small as 170 m/s you could move a 500 metric ton (mT) asteroid with just a single Centaur upper stage, at ca. 20 mT gross mass.

That's the case I want to look at because of the 5 to 10 year transfer time for the SEP case. It should be a shorter transfer time when using chemical propulsion. If you want, you can just calculate what would be the transfer time for the asteroid to get within the Moon's distance of the Earth. We can assume we use ballistic capture or small delta-v burn when it comes close to the Moon to put it in lunar orbit or at L2.

This sounds very interesting, if it can work. For reasons I don't fully understand, the Caltech study calls this a "prohibitive" mass cost.

Figure 19. The estimated propellant mass required to return a 1000-t NEA to lunar orbit would beprohibitive without solar electric propulsion (SEP).

ozz3jVI.png

Asteroid Retrieval Feasibility Study

http://www.kiss.caltech.edu/study/asteroid/asteroid_final_report.pdf

I'd note that's definitely outside the C3 ability of existing rockets; this would need an SLS launch at least, or in-orbit fueling.

Should also note that, if you're picking up an asteroid whole, the mass uncertainty means you can't feasibly exclude a 1,000 ton asteroid. (So your 500 ton figure understates it). But if you're going for a boulder option, then you could arrange that.

Also, don't forget there's a pretty large delta-v cost to match orbits with an NEA, even if it's extremely low in C3. That adds another fraction to the propellant ratio.

rIG4xCp.png

https://www.nasa.gov/sites/default/files/files/Creech_SLS_Deep_Space.pdf

Edited by cryogen

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The equilibrium temperature depends on the radiative properties of the spacecraft. Your temperatures are reasonable for planet surfaces, but for an artificial material with very low solar-emissivity, Teq can be much colder than -100 °C. Moreso, there's the trick that you can stack several layers of radiating surfaces, to get a combined, effective emissivity much lower than any physical material's. For a famous example, the James Webb telescope has a sunshield with five layers, εeff = 0.00038 (!!), and Teq = 40 K (-233 °C) on the cold side. That temperature is achieved solely by passive, radiative cooling.

https://i.imgur.com/3CpTttx.jpghttps://i.imgur.com/QMbeviY.jpg

http://jwst.nasa.gov/sunshield.html

In theory, you can get Teq so cold, that it's below the boiling point of liquid hydrogen (20 K at 1 atm). Alternatively, it can be slightly warmer, and the tank can be kept artificially cold by active refrigeration. Either way, no hydrogen is vented into space: it's "zero boil-off" (ZBO).

https://i.imgur.com/4va2TAH.png

http://develop.nttc.edu/sbipp/technologyportfolios/portfolios/ISS-Propellant_T-S/Archive%5C20060042833.pdf

- - - Updated - - -

This sounds very interesting, if it can work. For reasons I don't fully understand, the Caltech study calls this a "prohibitive" mass cost.

https://i.imgur.com/ozz3jVI.png

Asteroid Retrieval Feasibility Study

http://www.kiss.caltech.edu/study/asteroid/asteroid_final_report.pdf

I'd note that's definitely outside the C3 ability of existing rockets; this would need an SLS launch at least, or in-orbit fueling.

Should also note that, if you're picking up an asteroid whole, the mass uncertainty means you can't feasibly exclude a 1,000 ton asteroid. (So your 500 ton figure understates it). But if you're going for a boulder option, then you could arrange that.

Also, don't forget there's a pretty large delta-v cost to match orbits with an NEA, even if it's extremely low in C3. That adds another fraction to the propellant ratio.

https://i.imgur.com/rIG4xCp.png

https://www.nasa.gov/sites/default/files/files/Creech_SLS_Deep_Space.pdf

Thanks for responding. The idea that it would be doable by a Centaur for a 500 metric ton asteroid is based on the idea that the chemical propulsion transfer would be so short that it would be close to Earth both for the outbound and inbound flight. Then the delta-v to get there, perhaps by judicious, lunar gravity assists, would also be low, ca. 170 m/s. If so, then it would be a small proportion of the propellant load that would need to be burned to get there, and you would still be able to transport quite a sizable asteroid with the remaining propellant.

This is why it is so important to know whether it really is the case the travel time really would be short even with such a small delta-v of 170 m/s.

I agree with you that you would have to well characterize the size of the asteroid before you did the rendezvous.

Bob Clark

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The idea that it would be doable by a Centaur for a 500 metric ton asteroid is based on the idea that the chemical propulsion transfer would be so short that it would be close to Earth both for the outbound and inbound flight. Then the delta-v to get there, perhaps by judicious, lunar gravity assists, would also be low, ca. 170 m/s.

You left out the maneuver where you match velocities with the NEA, in sun orbit.

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