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Is it possible to use carbon monoxide as propellant for a rocket?

For example, carbon dioxide from the atmosphere of Mars could be transformed to carbon monoxide and oxygen via photochemical reduction using a photocatalytic process, and unlike the Sabatier reaction (producing methane), precious water is not consumed in the process.

Could the resulting products then be used as a potential bi-propellant?

$$2\,{\rm CO}_2 + 2h\nu\ \to\ 2\,{\rm CO} + {\rm O}_2$$

There may be issues of storage of LOX or LCO, but provided those are addressed somehow, could a rocket engine or thruster use these effectively?

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  • $\begingroup$ en.wikipedia.org/wiki/Photochemical_carbon_dioxide_reduction $\endgroup$ Apr 23 '17 at 10:00
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    $\begingroup$ This is a pretty interesting question actually! Split one oxygen from CO2 and you still have a gas (two in fact!), but split both and all you've got is a lump of coal unless you want to start spending your water. Could come in handy for some local applications, atmospheric recon craft for example. $\endgroup$
    – uhoh
    Apr 23 '17 at 12:26
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    $\begingroup$ As other gases, carbon monoxide could be liquified by low temperature. LCO (-191.5 °C) is colder than LOX (-183 °C) but not so cold as LH2 (-252°C) We should compare the specific impulse of this propellant with other combinations. $\endgroup$
    – Uwe
    Apr 23 '17 at 14:00
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    $\begingroup$ carbon monoxide is a less efficient combustible than methane but apparently will be cheaper to product on Mars $\endgroup$ Apr 23 '17 at 14:02
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    $\begingroup$ Here are two papers from NASA about this propellant combination: ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/19960045922.pdf ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/19910014990.pdf $\endgroup$
    – Uwe
    Apr 23 '17 at 14:23
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Yes, it could be. The ISP for CO/ O2 is about 200. Compare that to Methane, with a specific impulse of 299, and you can see it's really not that great.

Of some related interest is a hot Carbon Dioxide rocket, with a theoretical ISP of about 260. This would work for short surface to surface hops, but not beyond orbit.

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    $\begingroup$ Reading the paper in the report they reffer to ISP in the 260-290 range dependent on chamber preasure. Is the 200 ISP a typo? $\endgroup$
    – lijat
    May 29 '20 at 11:20
  • $\begingroup$ The quoted source supports your figures in exactly zero of the three numeric cases. You might want to re-read the quoted source, and edit your answer. $\endgroup$ Nov 16 '21 at 19:58
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LOX/CO propulsion systems will have a lower Isp compared to LOX/CH4 based systems. So for an equivalent delta-v (or mission application) this will translate to an increased amount of propellant that needs to be extracted from the Mars environment.

There is however an advantage to CO-based rockets on Mars: the propellant can be produced entirely from the atmosphere. This is a resource that will be easily accessible at any landing location.

The complete in-situ production of LOX/CH4 on Mars however will require accessing water in order to obtain the needed hydrogen in addition to atmosphere processing. Since liquid water cannot exist in a stable form on the surface of Mars, this will most likely require additional complex systems and hardware (i.e. landed mass) to extract buried water-ice deposits. Furthermore, the location of the surface mission will be dictated by where these water-ice deposits can be found.

So is it worth it?

The answer to this question likely depends on the mission application.

The CO-based rocket may appear preferable to the LOX/CH4 if the added propellant mass ends up being less than the equivalent system mass needed for water-ice ISRU extraction and processing.

For small robotic ascent vehicles this may be the case. However, for larger spacecraft (i.e. a human-scale Mars ascent vehicle with an inert mass of ~10,000 kg), the more efficient LOX/CH4 will likely quickly win the trade-off. This is especially true if the landed ISRU hardware is envisioned to be used for repeated missions.

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  • $\begingroup$ To date, there seems to be a lack of success in practical testing of CO/O2 engines. Apparently the combustion is hard to ignite and sustain, and the performance does strange things with changing chamber pressure. Solvable, of course, but not solved yet so it is rather hard to extrapolate actual performance and reliability figures. $\endgroup$ Nov 16 '21 at 20:02
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In TECHPORT_18280 published by nasa the summary talks about work on this. The teoretical ISP is aparently relativly high (324s) but the mixture apears to be hard to ignite in a safe way. The report then goes on about how such a safe ignition can be achived.

TECHPORT_18280 https://catalog.data.gov/dataset/o2-co-ignition-system-for-mars-sample-return-missions-phase-i

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