One year roundtrip, 30 day stay to Mars using solar thermal engines

[Frank]

One year roundtrip, 30 day stay to Mars.

Is it possible with solar thermal engines instead of nuclear thermal?

 

https://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/20140009587.pdf

We can assume Earth and Mars capture orbits for the solar thermal part and chemical rocket to and from capture orbits.

Perhaps with slower unmanned propellant drop missions ahead of mission if necessary.

 

 

 

Edited October 10, 2017 by Frank
oops

[Frank]

Enthalpy's threads say it can be done:

Solar Thermal Rocket

Manned Mars Mission

Non-Hohmann to Mars

 

Here's a  Delta-V map to get an idea of the scale of things:

https://upload.wikimedia.org/wikipedia/commons/8/8f/Delta_V_Earth_Moon_Mars.png

 

I like Enthalpy's idea of using STR (Solar Thermal Rocket) to spiral up from LEO.  Maybe it was implied, but the manned capsule would not be aboard for the slow trip, it would rendez-vous with it at, say EML-1 where the ships dock and maybe refuel.  Doing this would save a lot of launches compared to sending everything directly to EML-1.

 

It seems high delta-V is key for fast transit which is required to keep Radiation exposure minimal since shielding so far requires very large mass which is difficult to move.  This is where a STR or Nuclear Rocket or Electric Drive is important, though Musk's ITS also promises 80 to 150 days, but for a 2+ year stay using a Methane-Oxygen Rocket with 382s Isp. .

"The transport capacity of the 2016 spaceship from low Earth orbit to a Mars trajectory—with a trans-Mars trajectory insertion energy gain of 6 km/s (3.7 mi/s) and full propellant tanks—was projected to be 450 tonnes (500 tons) to Mars orbit, or 300 tonnes (330 tons) landed on the surface with retropropulsive landing.^[32] SpaceX estimated Earth-Mars transit times to vary between 80–150 days, depending on particular planetary alignments during the nine discrete 2020–2037 mission opportunities, assuming 6 km/s delta-v added at trans-Mars injection.^[32]"

https://en.wikipedia.org/wiki/Interplanetary_Transport_System#Passenger_spaceship

From reference 32 above, SpaceX expects to burn off 8.5 km/s in atmosphere at Mars and 12.5 km/s at Earth, so aerobraking instead of deceleration.

Given a hydrogen rocket Isp of 450s, higher speeds could be reached, and would require some deceleration, perhaps offset by the reduced mass of hydrogen?  Maybe ~15% faster?

Compare to the Nuclear mission in the previous post, SpaceX would take 90 days instead of 124 there and 206 back, though the nuclear short-stay mission is a greater distance...

 

Hmmm.

 

 

 

 

Edited October 11, 2017 by Frank
format

[Enthalpy]

I've re-read my documents, and the sunheat engine raises and lowers apoapses with kicks at periapses. This is more efficient than spiralling, but longer.

In my scenario, the manned vessels escape Earth or Mars by chemical propulsion (oxygen and hydrogen) to go fast, they accelerate by sunheat engines, and aerobrake at both destinations. So the transfer differs from presetting the return vessels and the descent-ascent modules on Martian orbit, which is slow. The manned vessel meets the preset hardware on Martian orbit.

I don't use any Lagrange points nor in-situ propellants. My scenario depends on no prior activity on Mars nor on landing at an accurate location, and offers redundancy.

Maybe other scenarios save some launch mass, but check the risks too.

If planning to use propellants produced on Mars or the Moon, consider burning hydrogen brought from Earth and producing only oxygen locally. It's almost as efficient as producing methane if the yield were 100%, and it's way simpler. No overcomplicated heavy plant brought to Mars. Electrolysis of dioxide in hot ceramic produces the oxygen, proven technology.
http://saposjoint.net/Forum/viewtopic.php?f=21&t=1953#p36418
obvious choice to my eyes.

----------

Hydrogen versus methane: the transit time improves very slowly with the delta-V and the specific impulse. It's better to save mass instead.

Aerobraking, sure. Even more so if the transfers are accelerated.

Transfer times: the figures by SpaceX are surprising. They may correspond to very favourable configurations of Earth and Mars. Anyway, don't compare them with other estimates that base on circular orbits. It would be wise to run a computation with SpaceX' figures, to check by how much they cheated optimized their example.

Please keep in mind that the announcement by SpaceX was part of a show. Nobody knows outside the company if they're serious about a Mars mission, nor whether the Raptor engine is meant for the BFR, or rather for a smaller launcher - nor how much soot plagues their gas generator at the Raptor. For Tesla's gigafactory, Musk had said "car batteries" but meanwhile he sells them for home and grid storage. About landing the Falcon 9's first stage on a barge, which is an overclever enabling solution, they didn't tell in advance, and once everyone saw it, they told "temporary solution".

Edited October 11, 2017 by Enthalpy

[Frank]

"Spiral" was not the right word, still learning the nomenclature, I just meant to say get to EML-1 efficiently and so, slowly, using solar engines.

 

I guess I'll try and lay out a first draft scenario...   Main ship assembled and fuelled at LEO, efficiently, slowly, using solar energy, reach EML-1.  Mate the upper stage with crew capsule of a Falcon Heavy fitted with a Raptor engine to the main ship which has all the habitat module, fuel tanks and solar thermal engines/concentrators.  Accelerate the ship with the Raptor engine.  Use the solar thermal engine to accelerate and decelerate to capture orbit at Mars.  Separate the upper stage and crew capsule for a Mars aerobraking landing.  Previous drops have fuel and life-support , habitat etc..  Orbital refuelling of the joined ship and similar return journey to Earth.  Separated main ship parks at EML-1 until the next mission.  Following the Musk plan for colonization, more ships would be added for each launch window.

No numbers yet or even a feasibility inkling...

 

I agree about hydrogen, I was thinking methane if long term storage is needed or hydrogen if usable quickly.  Boil-off seems to be an issue.

 

SpaceX numbers?  Don't know either way.  They may have a secret/alternate/improved/yet unknown  plan that differs from the original, would not surprise me.  I'm not keen on the drop 100 people off for 2+ years on the first try without even knowing if 38% gravity is enough to sustain life that long...

 

Why EML-1?  The idea of getting LOX from the lunar surface is appealing, if not in the near term, in the long term since getting Oxygen from the Lunar surface is less difficult than from Earth and can reuse upper stages of rockets as vehicles.  Lunar LOX could also be dropped at some Mars orbit and on Mars for refuelling.  This all assumes it is cheaper to get Oxygen from the moon than the extra cost of lifting it from Earth to space.

Also, a mass driver might ballistically launch fuel from Earth or Luna to EML-1 without much rocket Delta-V, basically just manoeuvring thrusters.

 

Raptor Upper stage?  That means 2..3.5 MN of thrust, perhaps convertible/augmentable with hydrogen instead of methane (it can be done in IC engines):

“Nevertheless, Raptor itself is clearly well on the way to full production, partly thanks to the US Air Force. Yup, SpaceX isn't the only organisation interested in this technology. From 2009 to 2015 Raptor development was funded solely by SpaceX, but in January 2016 SpaceX pocketed $33.6 million from the US Defense Department to develop a prototype version of an upper-stage variant of the Raptor designed to be used on the upper stage of a Falcon 9 and a Falcon Heavy.”

https://motherboard.vice.com/en_us/article/bmvpyw/how-spacexs-new-raptor-engine-will-get-us-to-mars

 

 

 

[Enthalpy]

Well, there are many possible options (plus the ones still not thought at), and even more combinations.

I doubt that propellants produced on the Moon and brought to Earth-Moon Lagrange point are any cheaper than if brought from Earth. And until I see a convincing proposal, I consider all cannons and variants are unusable.

You had initially the intention to evaluate a short stay mission. Its big delta-V constraints everything else, so you might begin with these figures and their implications on the propellant mass and so on. The speeds differ a lot from the Hohmann transfers illustrated in the second post. Please note that my xls supposes no flyby at Venus, but most scenarios seem to gain from it. And... I don't see any advantage in a short-stay mission, while the drawbacks are huge. Are you the last person putting time in it?

Methane and oxygen need active cooling to store in the Martian atmosphere, and I suppose in Earth and Mars orbit too. Once you have active cooling (a vital technology for space exploration, we should have had it for decades, wake up!), hydrogen is possible too; it's only a matter of tank mass and cooler difficulty. Here's a well insulated tank
http://www.scienceforums.net/topic/60359-extruded-rocket-structure/?do=findComment&comment=761740
But if sticking to dense fuels, some are about as efficient as methane, safe on Earth and won't freeze on Mars
http://www.chemicalforums.com/index.php?topic=56069.msg297847#msg297847
http://www.chemicalforums.com/index.php?topic=56069.msg272080#msg272080
(images only if logged in) while the toxic oxidizer Mon-30 is storable on Mars.

Aerobraking brings an awful lot, both at Mars and Earth. You must decide if aerobraking from the transfer speed or from orbit. If aerobraking from a fast transfer, you need wings to achieve downlift so the vessel stays in the atmosphere long enough to permit a bearable deceleration. It's probably not compatible with sunlight concentrators, so some choices are hard. But it has already been combined with a capture by Mars rather than landing. In my scenario, the crewed vessel aerobrakes, the preset modules don't. And you must decide whether the capsule or plane serves again at Earth.

If presetting hardware at Mars, you must decide whether on orbit or on the ground. Meeting on the ground suppose to land accurately, which is a risk, while orbiting hardware can be redundant and landed where and if needed. My opinion is very clear.

If producing propellants on Mars, you should check what power it takes over how long just to make oxygen. Ouch. Solar energy would need many big concentrators. And I consider methane is far too difficult and risky, while a little bit more hydrogen can be brought from Earth and consumed at the engines. Much safer, probably lighter than the methane plant.
http://saposjoint.net/Forum/viewtopic.php?f=21&t=1953#p36418

If meeting at Lagrange point (why? Isn't an elliptical Earth orbit better?) I'd bring much mass there by the craft with sunheat engine and little by the vessel with chemical engine: the astronauts, their radiation shield, their beds and suits, water and life support - but probably not propellants. And I have nothing against the Raptor, but it may be oversized for Earth escape, whose kick can last 600s: compute with the consumption rate. I found a few RL10 are strong enough, they burn the same hydrogen as the sunheat engine and offer a better ejection speed than methane.

A chemical engine is more efficient to escape Earth (it gains like 20-30%). I've evaluated the best share there
http://www.scienceforums.net/topic/76627-solar-thermal-rocket/?do=findComment&comment=818683
which you can compare with the sunheat engine alone there
http://www.scienceforums.net/topic/76627-solar-thermal-rocket/?do=findComment&comment=1009859
it gains 20-30%, interesting. The same holds at Mars! You choose if you use the same engines at Mars. And for a Hohmann Earth-Mars, the optimum is "nearly all' the transfer speed by the chemical engine, which translates to "all" since the optimum is wide, the sunheat engine only raising or sinking the apoapses. But an accelerated or a lowered perihelion transfer does use the sunheat engine during the transfer.

[Frank]

One thing about moon based Oxygen production is that it would be nice to know that ISRU works before we send stuff to a far-away planet.  In other words, if someone starts putting moon hotels in place (Bigelow) or decides that it's about time we have a moon base, then the Oxygen creation infrastructure must work.  Then sending fuel to EML-1 or Earth orbit or even Mars is possible and may even be economical.  I didn't begin to calculate the cost difference.  I'm just leaving the door open for that possibility, same with the mass drivers.  If these things ever get built and work as advertised, that's probably an order of magnitude cheaper at least.

 

" last person putting time in it? " - I don't follow.  In any case, I was under the impression the Solar Thermal Propulsion threads were about long stay and short trips, so I wondered if short stays and long trips were possible and how short can the trips can be given Solar Thermal Propulsion.

 

Well, methane can be stored in pressure tanks, but would need to be liquefied before use, as can oxygen.  Freezing is probably avoided by burying the tanks.  Hydrogen is tricky to get any kind of density as is evident from our lack of hydrogen cars, efficient storage is a problem.  Active cooling is a possibility - sure.  I mostly went with methane based on the Raptor engine and the idea that fuel can be synthesized on Mars (someday).  I'm not a chemist, so I assume smart people are working for Musk.

 

Aerobraking I took from the Musk plan.  Assuming it would work for a smaller craft.  Now that I think about it, it isn't clear how the dragon capsule and the glider would work together.  I guess they don't which is too bad, because I liked the escape feature.  Maybe a fairing style of glider skin with a spare one on the main ship for the Earth reentry.  Needs more work...  As for re-use, returning upper stages to Earth seems too expensive still, so my thought there was to re-use the upper stages in space.  Even though the upper stage used for the Mars mission would return to Earth it seems that the 5 refills and two re-entries would be enough re-use .  Maybe burn up on Earth re-entry with a capsule landing instead of a glider landing?

 

Accurate landings on Earth have been going very well for SpaceX hitting the bull's eye 18 times in a row now.  I had though both orbit and land drops, no ISRU at first, though setting up the equipment for future missions might be a part of the short stay mission.  The main idea for me is to re-fuel at Mars and make the manned trips as short as possible.

 

The important thing about the Raptor vs the Merlin was the fuel type as RP-1 isn't going to help anyone.  3 MN would sustain gravity acceleration for a 300 ton ship with a single engine!  Artificial gravity anyone?  It is also more efficient.  Another thought is that having at least two engines would be nice redundancy should one fail on landing for example.  They did develop a scaled version of the Raptor that might work better for this application.  Hydrogen vs Methane - still not convinced either way.  Ammonia or Ammonia hydrogen-capable might be good too.  Don't know.

 

If you're saying that chemical rockets are more efficient than solar thermal for LEO to staging area (EML-1 or high orbit), then why bother with staging at LEO at all?  Just go straight there, robotically assemble/refuel at EML-1.

 

[Enthalpy]

On Mars, oxygen would be produced from the CO₂ atmosphere, for the crew and maybe the engines. Demonstrators exist already that electrolyse  CO₂ in a hot ceramic. Trials on the Moon would be too different.

Short stay: your choice, your good right.

Methane can't be liquid above 191K, whatever the pressure. It demands active cooling on the too warm Mars. Storing gases is excluded because the tanks are too heavy for space travel. Even if burying in the chilly Martian soil, some fuels like RP-1 "kerosene" would freeze; concentrating sunlight is a better option, and low-freezing fuels a safer one.

The density of rocket propellants isn't a primary concern. Hydrogen serves on most launchers, even at the first stage on Delta.

Produce methane on Mars: smart people for sure, but nevertheless I say it's a bad idea. Very difficult, and bringing a bit more hydrogen is much simpler, safer, probably lighter. Remember hydrogen must be brought and stored to produce methane anyway.

Glider and escape: have a look at the IXV. Put it atop a launcher without fairing but with an escape rocket on top.

Accurate landing: a mission to Mars was lost recently because of inaccuracy. 18 successes on Earth prove less than 95% reliability for that single failure cause. "We landed the crew too far from the return vessel and can't send them anything timely" would be disastrous. Zubrin's team foresees a rover but neglects the lack of highways and the insufficient radiation shield.

Other companies exist outside SpaceX, with engines for varied fuels. The RD-170 pushes 8MN with RG-1. Methane is negligibly more efficient but much more dangerous, while RP-1 doesn't catch fire with a lighter. And please forget ammonia, it is t-o-x-i-c, didn't you know?

"If you're saying that chemical rockets are more efficient than solar thermal for LEO to staging area": No, I didn't.

[Frank]

Hydrogen has lots of handling issues, maybe they are all solved, but I suspect that the SpaceX team did the cost comparison and decided to go with methane.

True, the Lunar ISRU is different, but experience from a Lunar settlement would still be helpful.  I don't feel it's an either-or case anymore with private money going into space exploration.  Both Lunar outposts AND Mars outposts/colonies are possible concurrently.  SpaceX mentioned the Sabatier reaction  https://en.wikipedia.org/wiki/Sabatier_reaction to convert CO₂ + 4H₂ → CH₄ + 2H₂O   ∆H = −165.0 kJ/mol, so Mars CO2 with Earth hydrogen to get methane and water.  I'm guessing the boil-off rate is manageable to feed the reactor?

Carbon-fibre tank vs hydrogen tank plus insulation..

Energy density of hydrogen 8.491 MJ/L, liquid methane (LNG) 22.2 MJ/L  compressed methane (CNG 250 bar) 9 MJ/L.  So both tanks are the same size.  Mass difference?

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

Ammonia can be handled safely, it's used on crops everywhere, regardless, it has lower energy content than methane, so not ideal as rocket fuel, though it might combine well with thermal thrusters.  A space suit is probably enough protection and must be worn anyway.

Well, that lander crashed, so there was no recovering.  A rover is one solution.  A Mars short hop is another if enough fuel is left in the rocket - this would assume a navigation failure.  Risky business for sure.  A short hop using one of the rockets previously landed and partly refuelled is another.

 

 

 

 

[Frank]

So, it seems using Solar Thermal Propulsion for fast transit to Mars isn't helpful.  Aerobraking a chemical rocket (the SpaceX way) is better.

STP is probably useful to bring payloads to higher orbits or to EML-1 or to low lunar orbit maybe.

STP is probably useful to haul (non-human) payloads to Mars and back, as well as hauling from Mars lower orbit to a higher orbit.

 

I used this spreadsheet to get there https://space.stackexchange.com/questions/18396/how-is-80-to-150-days-to-mars-possible-at-6-km-s-tmi-delta-v and here http://clowder.net/hop/railroad/Hohmann.xls

 

This work on STP is interesting: https://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/20160003173.pdf

 

 

[Enthalpy]

On 10/15/2017 at 9:29 PM, Frank said:

Hydrogen has lots of handling issues, maybe they are all solved, but I suspect that the SpaceX team did the cost comparison and decided to go with methane.

True, the Lunar ISRU is different, but experience from a Lunar settlement would still be helpful.  I don't feel it's an either-or case anymore with private money going into space exploration.  Both Lunar outposts AND Mars outposts/colonies are possible concurrently.  SpaceX mentioned the Sabatier reaction  https://en.wikipedia.org/wiki/Sabatier_reaction to convert CO₂ + 4H₂ → CH₄ + 2H₂O   ∆H = −165.0 kJ/mol, so Mars CO2 with Earth hydrogen to get methane and water.  I'm guessing the boil-off rate is manageable to feed the reactor?

Carbon-fibre tank vs hydrogen tank plus insulation..

Energy density of hydrogen 8.491 MJ/L, liquid methane (LNG) 22.2 MJ/L  compressed methane (CNG 250 bar) 9 MJ/L.  So both tanks are the same size.  Mass difference?

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

Ammonia can be handled safely, it's used on crops everywhere, regardless, it has lower energy content than methane, so not ideal as rocket fuel, though it might combine well with thermal thrusters.  A space suit is probably enough protection and must be worn anyway.

Well, that lander crashed, so there was no recovering.  A rover is one solution.  A Mars short hop is another if enough fuel is left in the rocket - this would assume a navigation failure.  Risky business for sure.  A short hop using one of the rockets previously landed and partly refuelled is another.

 

 

 

 

The Sabatier reaction is known, yes. Engineering is an other story. As soon as 50% of the hydrogen doesn't make methane, the conversion to methane needs more hydrogen than if you burn it directly in the rocket engine. And then you must transport the plant there and operate it, plus land the crew at the proper location.
Once again, "they have their reasons" is not a receivable argument for me.

Guesses are nothing. How do you achieve it, how heavy is it?

Your "energy density" is just irrelevant. You should re-consider the mass of a gas tank. Put figures on it. Without figures, there is no science nor engineering.

You mistake liquid ammonia with an ammonia solution. And on crop, it kills regularly. Putting 10t on 100t in a launcher would be insane.

Refuelled: how? "Short hop": how far, how much propellants, obtained how?

On 10/18/2017 at 3:17 AM, Frank said:

So, it seems using Solar Thermal Propulsion for fast transit to Mars isn't helpful.  Aerobraking a chemical rocket (the SpaceX way) is better.
STP is probably useful to bring payloads to higher orbits or to EML-1 or to low lunar orbit maybe.
STP is probably useful to haul (non-human) payloads to Mars and back, as well as hauling from Mars lower orbit to a higher orbit.

How do you reach such bizarre conclusions? "Seems, probably, probably". What script, delta-V, mass?

The whole attempt makes me uneasy. The general impression is that you have less expertise on the topic than would be necessary and hope to compensate it by software found on the Web and by compiling documents. On a project that has never been done before, assembling existing data doesn't bring a solution.

==========

You might want to explore the possibility of bringing water from a main-belt comet to Martian orbit and electrolyse it there to make propellants for a descent-ascent module and for a return module. I checked there how much can be transported
http://www.scienceforums.net/topic/76627-solar-thermal-rocket/?do=findComment&comment=757109
addenda and corrections followed
http://www.scienceforums.net/topic/76627-solar-thermal-rocket/?do=findComment&comment=757663
I didn't put figures on such a Mars trip.

and in case you still consider the opposition mission, you could evaluate if this radiation shield brings something
http://www.scienceforums.net/topic/80982-shield-astronauts/
for which you'd have to evaluate the improvement on Solar radiation and check if the extrasolar one is acceptable.

[Frank]

1 hour ago, Enthalpy said:

The Sabatier reaction is known, yes. Engineering is an other story. As soon as 50% of the hydrogen doesn't make methane, the conversion to methane needs more hydrogen than if you burn it directly in the rocket engine. And then you must transport the plant there and operate it, plus land the crew at the proper location.
Once again, "they have their reasons" is not a receivable argument for me.

Guesses are nothing. How do you achieve it, how heavy is it?

Your "energy density" is just irrelevant. You should re-consider the mass of a gas tank. Put figures on it. Without figures, there is no science nor engineering.

You mistake liquid ammonia with an ammonia solution. And on crop, it kills regularly. Putting 10t on 100t in a launcher would be insane.

Refuelled: how? "Short hop": how far, how much propellants, obtained how?

How do you reach such bizarre conclusions? "Seems, probably, probably". What script, delta-V, mass?

The whole attempt makes me uneasy. The general impression is that you have less expertise on the topic than would be necessary and hope to compensate it by software found on the Web and by compiling documents. On a project that has never been done before, assembling existing data doesn't bring a solution.

==========

You might want to explore the possibility of bringing water from a main-belt comet to Martian orbit and electrolyse it there to make propellants for a descent-ascent module and for a return module. I checked there how much can be transported
http://www.scienceforums.net/topic/76627-solar-thermal-rocket/?do=findComment&comment=757109
addenda and corrections followed
http://www.scienceforums.net/topic/76627-solar-thermal-rocket/?do=findComment&comment=757663
I didn't put figures on such a Mars trip.

and in case you still consider the opposition mission, you could evaluate if this radiation shield brings something
http://www.scienceforums.net/topic/80982-shield-astronauts/
for which you'd have to evaluate the improvement on Solar radiation and check if the extrasolar one is acceptable.

Those are all good points.

You may also be correct that the Musk/SpaceX plan is wrong and ill-conceived.  But the biggest trick is money, no money, no space travel.  I don't know how to compare the cost of a cryogenic hydrogen tank VS the cost of a pressurized methane tank.  I'm not the one trying to prove SpaceX is wrong.  My risk is thinking they can do what they claim and I may be dismissing a better concept plan than their concept plan.

 

Anhydrous Ammonia is conveyed by train all over the place in tankers that are many tens of tons each and they do crash and rupture (sadly).  Fuels are dangerous, some are poison, some explosive.  Risk assessment must be done.  I wasn't suggesting it be used blindly, just not rejected out of hand because of toxicity.

 

The SpaceX plan either has a rocket fuelled and ready to launch remotely, or people are there refuelling a rocket from ISRU or maybe it's all meant to be done robotically.  In any case, the plan is that there is a rocket (or many) sitting there from a previous supply/cargo mission.  There would need to be a way to transfer fuel from a cargo rocket to a crew rocket.  The hop with another rocket was MY idea addressing your concern about accuracy, SpaceX has only said they would land on target.

 

What I found was that the time it takes to accelerate a ship using STP can be 30 or more days and that the distance travelled in that time is greatly inferior compared to a rocket burn, so that higher speed must be attained to compensate for the penalty.  When greater speeds are attained, the extra arrival speed must be burned off to reach a capture orbit which also takes more time compared to a rocket burn or direct aerobraking.  There MAY be a way to do it, I haven't found one that is significantly faster than 80 days, so I'm putting it aside.  I did make a spreadsheet of my own, put haven't posted it.

 

"probably useful" means that STP is a way to do it, but it might be cheaper to do a one-way trip with a rocket than a two way trip with STP or there are other, better ways to do each of the missions.   I haven't done the math yet, or looked at all the possibilities (who can?) someone who's already done it can post the results or I may do it in the future. 

 

Yes, I'm just starting to learn this stuff.   It's interesting and I've learned a lot already.

 

[Enthalpy]

SpaceX: nobody know what their plans are. What they tell at AIAA meetings and Press conference is designed to make buzz. You don't even know if they plan to go to Mars. So you should not deduce any technical choice based on their claims.

More generally, science and technology can't work on rumours, assumptions, impressions, interpretations of supposed reasoning by other people. It boils down to "how". You decide how you realise a function and convince with shapes, figures... that it's realistic. Anything else would fit in an annual conference of space frenzy, but would not be science nor engineering.

Here you accumulate bizarre deductions on wrong hypotheses based on dubious claims. The result isn't convincing.

I've already provided figures about the travel time to Mars using the sunheat engine. Why add your impressions to this and lead to wrong conclusions? Check my figures, and if they're wrong, tell us where and why.

Ah, and I'd prefer fewer acronyms. I know that my fellow engineers and scientists love to use and misuse them everywhere. It could be by imitation, to look like a true one, but it's very bad practice.

[Frank]

1 hour ago, Enthalpy said:

I've already provided figures about the travel time to Mars using the sunheat engine. Why add your impressions to this and lead to wrong conclusions? Check my figures, and if they're wrong, tell us where and why.

I didn't say a short-stay mission can't be done with solar thermal, I wrote "So, it seems using Solar Thermal Propulsion for fast transit to Mars isn't helpful.  Aerobraking a chemical rocket (the SpaceX way) is better. " 

I like simplicity and to me, using a chemical rocket (which we know works) and aerobraking (which we know works) to do a mission seems better than prototyping a solar thermal engine that will allow a spacecraft that can do aerobraking to be pushed out to Mars.  Are there huge benefits to using STP I should consider for a short stay mission?