Reusable Mars Ascent/Landing Vehicles

[Kaputnik]

Many scenarios involving significant human settlement of Mars envisage hypothetical reusable vehicles capable of reaching Mars orbit and of landing safely on the surface again, using in-situ propellants, most commonly LOX and CH4.

I wanted to start this thread to discuss the potential development of such a vehicle. I think it would be safe to assume that the vehicle would source its propellants externally, from a dedicated ISRU plant, but what would be the best propellant choice? Should it be single stage or multiple stages? Should it use primarily a propulsive landing system or would it be better to have a lifting body or reusable capsule design?

[Dalhousie]

Simplicity suggests single stage to orbit.

Propellants - methane-LOX, ethylene-LOX, methanol-LOX, CO-LOX would be my first guess.

A quick search found the following studies:

MSR missions

https://e-reports-ext.llnl.gov/pdf/235739.pdf

http://www.sciencedirect.com/science/article/pii/S0094576599001484

Crewed missions http://www.nss.org/settlement/mars/AccessToMars.pdf

http://www.lpi.usra.edu/publications/reports/CB-1063/Rowan.pdf


The biggest challenge might be the servicing and turn round between flights.

[go4mars]

I wanted to start this thread to discuss the potential development of such a vehicle.

Starting with a program like lunar lander challenge where the required burn is similar to expected requirements on Mars (let's arbitrarily call it "Grasshopper" test program) would be useful.  Whether or not "Grasshopper" has this as its primary goal or not, it is admittedly useful for this purpose as well. 

I think it would be safe to assume that the vehicle would source its propellants externally, from a dedicated ISRU plant, but what would be the best propellant choice? Should it be single stage or multiple stages?

Using methane/oxygen has some advantages: Means the folks on Mars would have a necessarily robust & plentiful source of water and oxygen,  isp beats CO, thrust better than argon systems. 

Should it be single stage or multiple stages?

Single.  A Grasshopper should be able to put a dragon into Mars orbit and re-land.  Potentially with power to spare. 

Should it use primarily a propulsive landing system or would it be better to have a lifting body or reusable capsule design?

As long as it can land just about everywhere on the planet for exploration or to fuel up at ice locations, then anything goes. 

[Kaputnik]

Thanks, very useful links.
Supersonic Retropropulsion seems to be the key enabling technology.
I see in the NSS paper they assume an isp of 465s which suggest a regen engine akin to the RL10 or RL60. During the initial entry period the engines, which are rigidly fixed to the heatshield and exposed at all times, are not running. This surely introduces problems with heat rejection, to which I do not see an obvious answer.

Perhaps one method of getting around the problem is to have dedicated high-isp ascent engines placed centrally, with retractable protective covers, whilst descent is handled by separate engines located on the sidewall (like Dragon) to minimise heating- or alternatively, the descent engines could use a different cooling system allowing them to be exposed to the airflow.

[Kaputnik]

Starting with a program like lunar lander challenge where the required burn is similar to expected requirements on Mars (let's arbitrarily call it "Grasshopper" test program) would be useful.  Whether or not "Grasshopper" has this as its primary goal or not, it is admittedly useful for this purpose as well.

Erm, that could cause a lot of confusion. Grasshopper is almost certainly not anything to do with landing on Mars. Why not use a different name?

Using methane/oxygen has some advantages: Means the folks on Mars would have a necessarily robust & plentiful source of water and oxygen,  isp beats CO, thrust better than argon systems.

I was really meaning LOX plus a choice of LH2, CH4, or more complex hydrocarbons. CO seems unlikely to have the necessary isp. Electric propulsion using argon is for a completely different application so not in contention, in the same way that a solar sail is not.

Should it be single stage or multiple stages?

Single.  A Grasshopper should be able to put a dragon into Mars orbit and re-land.  Potentially with power to spare.

I wasn't meaning 'should' as in 'what would you like to see in an ideal world'. More a case of asking what is the best engineering compromise.
For example, a two stage vehicle could have a first stage that uses fully propulsive landing, allowing the orbital stage to be optimised for the ascent, rendezvous, and re-entry phases.
Also, your description of a Grasshopper+Dragon is a two stage vehicle when viewed as a complete system for reaching and returning from orbit.

[Robotbeat]

Any propellant combination OTHER than CO/LOx has serious logistics and/or infrastructure limitations. CO/LOx (at near vacuum) can reach ~300s isp, higher than almost any propellant combination at Earth sea level other than hydrogen. CO/LOx is also pretty dense.

I understand the advantages of using methane instead, but I think you are down-playing the difficulty of methane ISRU versus CO ISRU.

[e of pi]

I understand the advantages of using methane instead, but I think you are down-playing the difficulty of methane ISRU versus CO ISRU.

Particularly if this is to operate at Mars for multiple flights, since methalox would require a source of seed hydrogen, while colox just needs the martian atmosphere and a power source.

[Kaputnik]

I would argue that we are talking about a resuable vehicle operating at a large, permanently manned base- not just the initial missions. In this context, I think it is entirely reasonable to assume that an advanced ISRU capability could be in place at this time, including the extraction of water from subsurface sources.
The NSS paper suggests LOX/LH2 with an isp of 465s- this is a big increase over LOC/CO.

[Robotbeat]

I would argue that we are talking about a resuable vehicle operating at a large, permanently manned base- not just the initial missions. In this context, I think it is entirely reasonable to assume that an advanced ISRU capability could be in place at this time, including the extraction of water from subsurface sources.
The NSS paper suggests LOX/LH2 with an isp of 465s- this is a big increase over LOC/CO.

What about density? You haven't mentioned that. It's just as important, because if it's (say) 10 times easier to make CO/LOx versus LH2/LOx and the dry mass is the same or slightly less for CO/LOx, then it still makes sense to go with CO/LOx.

I agree that the ideal propulsive method may well change as infrastructure is built, but for the initial stages, CO/LOx is a lot simpler with a lot fewer infrastructure requirements. CO/LOx ISRU can be done anywhere on the planet and without needing any Earth-moving equipment at all, thus it's more flexible. You could set up small, automated ISRU production facilities at strategic locations on the planet, allowing a lot more flexibility and capability for the same level of infrastructure.

The higher Isp of hydrolox is more important at Earth, where the infrastructure to produce hydrogen is more readily available and where the delta-v to orbit is a lot higher. For small delta-vs where the exponential nature of the rocket equation doesn't come into force as much, the penalty from lower Isp is a heck of a lot less and the volume advantage of a denser fuel becomes more important.

[go4mars]

Grasshopper is almost certainly not anything to do with landing on Mars.

Your assertion hasn't been well established. 

your description of a Grasshopper+Dragon is a two stage vehicle when viewed as a complete system for reaching and returning from orbit.

I think dual use might be the better term (versus 2-stage).  There could be places inside or on top of the "grasshopper", where ice, iron, random stuff could be stored.  It could fly independent of grasshopper.  But if people go in it, dragon (2), affixed to the top, would provide handy LAS capability, redundant life support, etc.  Wouldn't necessarily even need to separate from grasshopper at any stage in point-to point travel. 

Any propellant combination OTHER than CO/LOx has serious logistics and/or infrastructure limitations. CO/LOx (at near vacuum) can reach ~300s isp, higher than almost any propellant combination at Earth sea level other than hydrogen. CO/LOx is also pretty dense.

I understand the advantages of using methane instead, but I think you are down-playing the difficulty of methane ISRU versus CO ISRU.

Namely, finding and melting ice...  Ice seems to be pretty widespread.  Perhaps there would be 3 tanks on this thing.  CH4, O, and H.  some low boil-off H could be stored on board perhaps from an initial trip to the Martian poles?  Or perhaps more likely, as Ice.   H storage might even take the form of high density packing on kerogen with high sorptive capacity (since it doesn't need to be drawn down quickly).           

[RocketmanUS]

CO, how stable is it to store for long periods of time?

What is the fuel mixture for a CO/O2 rocket engine?

[A_M_Swallow]

The other thing to remember are pumps and pressurisation.  CO₂ and Argon are ISRU gasses available on Mars.

[Kaputnik]

What about density?

I agree density is important. The trades might be interesting on a Mars ascent-descent vehicle- low density could actually be necessary or useful in creating a viable entry/landing vehicle.

I agree that the ideal propulsive method may well change as infrastructure is built, but for the initial stages, CO/LOx is a lot simpler with a lot fewer infrastructure requirements.

If a viable vehicle can be created using CO/LOX then that would be good news. However IMHO a reusable vehicle is unlikely to feature in 'initial' stages of infrastructure buildup. This is a long term idea.

The higher Isp of hydrolox is more important at Earth... where the delta-v to orbit is a lot higher.

Agreed, but I am talking specifically about a reusable surface to orbit to surface vehicle. It has a total delta-v considerably higher than a pure ascent vehicle. It's design is further complicated by the aerodynamic problems of entry. All in all, it would be a very challenging vehicle to create, needing very good performance whilst being rugged enough to be field servicable.

[Kaputnik]

Grasshopper is almost certainly not anything to do with landing on Mars.

Your assertion hasn't been well established.

I think the onus is on you to prove that Grasshopper is anything other than a 1st stage propulsive recovery testbed.

Wouldn't necessarily even need to separate from grasshopper at any stage in point-to point travel. 

OT. This thread is about fully reusable orbital vehicles, not PTP.

[Kaputnik]

The other thing to remember are pumps and pressurisation.  CO₂ and Argon are ISRU gasses available on Mars.

Good point. Lack of He might make a LH2 based architecture very difficult or even impossible. Can Ar be used instead? What about Ar or N for use in pressurising and purging tanks fed by the other fuel options?