New Mars Forums

I ran some numbers on another forum about replacing scuba tanks.  It should be doable technically to do electrolysis to separate the oxygen out of water at a much lighter weight than a scuba tank, which run about 15 kilos for about an hour's worth of breathing time underwater.

To chemically separate the hydrogen and oxygen in water takes about 16 million joules of energy per kilo of water. This is at 100% efficiency. I don't know what the best efficiency actually is now but we'll see the amount that can done be using batteries is so high that likely it can be better than using scuba tanks for the same weight.

The ratio of oxygen to hydrogen in water by mass is 8 to 1. So using 16 megajoules of energy you get 8/9 of a kilo of oxygen. This means 16*(9/8) = 18 megajoules gives 1 kg of oxygen. Using as an estimate that oxygen usage for astronauts would be similar to scuba divers we need .5 kg of O2 for 7 hours underwater, or 1 kg of O2 for 14 hours.

Now for the weight of batteries. See this table for a list of energy densities per weight:

https://en.wikipedia.org/wiki/Energy_de … materialsy

The best energy density among the commonly available batteries is the lithium battery (non-rechargeable) at 1.8 megajoules per kilo. Then to provide the 18 megajoules to get 1 kilo of oxygen you would need 18/1.8 = 10 kilos of the lithium batteries. I'll assume the weight of the system will be dominated by the weight of the batteries since electrolysis can be done simply by placing electrified wires in the water. So at 15 kilos of the batteries, the same weight as the scuba tank for 1 hour of breathing time, you would actually get enough oxygen for (15/10)*14 = 21 hours. Or said another way, to get enough oxygen for 1 hour would only take 10/14 = .7 kilos, about 1.6 pounds.
(Note: Edited calculations.)

Practical problems of course are for one how efficient is the electrolysis procedure? Also I believe the lithium batteries catch fire when wet. You would need to insure the battery casings are super-waterproof.

Then there is the weight of the electrolysis components, and the weight of the tank to hold the oxygen as it is being produced. You would also need to dispense with the hydrogen in a way to insure it does not combine with the oxygen to catch fire. It might be sufficient for this to mix it with surrounding water.

Though this would be technically doable the biggest problem is that scuba tanks do not use pure oxygen because it is toxic at depth. Perhaps GW can answer a question on this. The ISS astronauts would not need to prebreathe if their suits were at .8 bar O2. If O2 was given to divers at this pressure could it be used at depth without toxic effects?

This method could also be used to replace the heavy compressed oxygen tanks in astronauts suits by just carrying the small amount of water instead. The Primary Life Support System used on the astronaut spacesuits weighs 38 kilos. The power to separate out the approximate .54 kg oxygen out of water would only take about 5 kg of lithium batteries.

However, for the Mars spacesuit case, it might be more advantageous to separate out the O2 from the CO2 atmosphere.  How much power is needed to separate 1 kg of O2 out of CO2? If the power required is not too high you may even be able to do it with solar cells so the astronauts could spend virtually unlimited amounts of time on the surface in their suits.

  Bob Clark

It occurs to me a bigger energy cost will be involved in increasing the tiny oxygen partial pressure to that used in NASA spacesuits, ca. 0.3 bar. I'll look up some formulas for that.

  Bob Clark

Thanks. Pretty cool.

   Bob Clark

The New Horizons mission to Pluto further confirms the great interest the public has for planetary missions. The Mars Curiosity mission got over a billion hits to the NASA web page over a year. And New Horizons mission web page got 10 million hits on July 14th alone.

A lander mission to Europa to explore the subsurface ocean could result in the most revolutionary scientific discovery in human history: the discovery of life on another world. Such a discovery would dwarf even the Apollo missions in importance.

Then it is notable that following the commercial space approach such a mission could be privately financed at the few hundred million dollars range. This would be low enough considering the great interest such a mission would get that it could even be profitable from advertising.

Then I have changed my sig file to indicate such a revolutionary mission could be accomplished at such low cost and in a near time frame.

Bob Clark

The title is misleading since a NASA official is quoted as saying the station keeps a six-month leeway in needing supplies.

  Bob Clark

After a calculation I was surprised to discover that with orbital propellant depots, manned interplanetary flights become scarcely more expensive than flights to LEO. That is, spaceflight to the Moon, Mars, Venus, asteroids could become as routine as flights to the ISS are now just by having propellant depots in place at departure and arrival points.

For instance rather than the 1,000 metric tons(mT) estimated needed to be launched to orbit for a manned Mars mission, a single empty Falcon 9 first stage at ca. 15 mT dry mass could do ALL the propulsion stages by itself, when you have propellant depots at Earth orbit and Mars orbit. No huge, and hugely expensive, Mars Colonial Transport, SLS, or even Falcon Heavy required.

And it's not just the Falcon 9 first stage. It would also work for the first stages for the Atlas V, the Delta IV, the Ariane 5, the Soyuz, etc., with the empty stages lofted to orbit and refueled in orbit at both departure and arrival points. This means every space faring nation could do Mars missions for little more than currently used to send flights to the ISS.

Here is the calculation the Falcon 9 first stage could to the round trip flight with orbital refueling:

The Coming SSTO's: Applications to interplanetary flight.
http://exoscientist.blogspot.com/2012/0 … ns-to.html

This was focusing on single-stage-to-orbit (SSTO) vehicles but this is not a requirement. Any currently in use medium size launcher's first stage would work. You would need low cost propellant delivered to orbit. Fully reusable launchers bringing the price to orbit down to ca. $100 per kilo would do it.

But it could also be done with the propellant obtained off-Earth. Some Mars advocates have been opposed to propellant depots since it was thought the propellant had to be obtained from the Moon, requiring giga-dollar expenditures. But it doesn't have to be from the Moon. For instance there are near Earth comets that experience outgassing, as has been observed by the Rosetta spacecraft:

Dust Whirls, Swirls and Twirls at Rosetta’s Comet.

by BOB KING on MARCH 9, 2015
http://www.universetoday.com/119296/dus … tas-comet/

In such a case, no landing or mining would be required. You would just collect the released H2O, and CO2, CO for hydrocarbon fuel, from orbit.

  Bob Clark

Thanks for those links. I'm in favor of fast travel times to Mars. Rather than using propulsion to slow down, aerocapture would allow a smaller size vehicle to be used, and cheaper mission cost.

   Bob Clark

Interesting report here:

Solar Electric Propulsion.
Technology Development.
www.nasa.gov/sites/default/files/files/CTaylor_SEP.pdf

It discusses the idea of using chemical propulsion for getting out of Earth's gravity well then using SEP to reach a BEO destination. Note this is how most interplanetary probes that used electric propulsion worked. The report has this surprising chart near the end:

It includes the interesting fact that using the same propulsion system at 50 kW power, a 2,000 kg dry mass spacecraft would take 91 days to make the flight to Mars when running at 4,000 s Isp, but only 23 days(!) when running at 2,000 s Isp. This make sense because the thrust drops off for EP thrusters as you increase the Isp. But the thing is we already have a 50 kW Hall effect thruster in the NASA-457m:

Performance Test Results of the NASA-457M v2
Hall Thruster.
http://ntrs.nasa.gov/archive/nasa/casi. … 014613.pdf

I was puzzled at first by the graphic on the side listing delta-v's to reach various interplanetary destinations. For Mars, it's listed as 5.6 km/s. Since the report is discussing using EP after leaving Earth's vicinity, this must mean the delta-v just to leave Earth's position in the Solar System to Mars position, so not including the delta-v for escape velocity or slowdown delta-v at Mars.

Support for this is suggested by this:

Hohmann transfer orbit

http://en.wikipedia.org/wiki/Hohmann_tr … t_transfer

  Bob Clark

Thanks for those calculations. As you can see from your calculated values, a major problem with the VASIMR is the mass of the engine. You cited the report:

The VASIMR Engine: Project Status and Recent Accomplishments.
AIAA 2004-0149

http://spaceflight.nasa.gov/shuttle/sup … ss2004.pdf

So with the numbers used in this report of 1 megawatt (input) power and weight to power ratio of 1 kg to 1 kW power, the engine masses in this report at ca. 1,000 kg. Then this graph indicates a thrust-to-weight ratio in the range of 1 to 1,000, or perhaps 1 to 500 if using lithium, which they say is not the preferred propellant because of difficulties of handling it.

However, already existing Hall effect thrusters can get thrust-to-weight ratios of 1 to 100:

Developmental Status of a 100-kW Class
Laboratory Nested channel Hall Thruster.
IEPC-2011-246
Table 1, Example of concentrically NHT specific mass and footprint savings, p. 5.

http://pepl.engin.umich.edu/pdf/IEPC-2011-246.pdf

Remember being able to get high thrust is important for this scenario of fast travel time. Otherwise it's the same scenario as with ion propulsion where the travel time would actually be slower than chemical propulsion because of the very small thrust.

Here are the numbers you are envisioning for your VASIMR vehicle:

But using Hall effect thrusters you could reduce the calculated weight of 17 tonnes to 1.7 to 3.4 tonnes.

The next problem is the mass of the solar power system. Based on a 300 watt/kg specific power for the solar cells you get a 20 tonne mass for the solar power source. However, I think at least 1,000 watt/kg solar power is possible by using solar concentrators. The reason is the mirrors or lenses would be much lighter than solar cells used so you could get the same power at lighter weight.

So if the solar power sources mass, say, 7 tonnes then you see the case may indeed close for getting the total vehicle dry mass low enough to get the high speed needed, but by using existing Hall effect thrusters, not VASIMR.

  Bob Clark

Thanks for that insider info about engineering. I think I remember SpaceX presenting proudly how young the average age of their engineering staff was. Maybe. But they still need some of the "old guard" to transfer the art to the new generation.

   Bob Clark

Perhaps H2 could be used for the fill gas for a ballute. How much total energy would need to be dissipated by a heat shield on reentry?

  Bob Clark

We're you able to hear the latest full scale test of the F9 first stage:

http://www.nasaspaceflight.com/2015/02/ … ad-launch/

  Bob Clark