NASA adopts the best of Zubrin's ideas, eventually

[QuantumG]

Once upon a time, it seemed that Zubrin's name was unspeakable at NASA.. and you can understand why.

They used to dismiss the leverage of Mars ISRU, but now it's a standard part of any architecture. Same with promoting conjunction class missions over opposition class, to reduce total radiation exposure times.

I think Zubrin's next contribution will be dismissing the myth that crew safety can be improved with hyperbolic transfer trajectories (aka VASIMR pixie dust), as no-one has been doing cost-benefit analysis of the reduced radiation exposure to the elimination of abort modes.

Another good point he's made lately is to remind us that we simply don't know what prolonged exposure to galactic cosmic radiation does to the human body. We have hunches, and assumptions, but no hard data. Human factors experts and flight surgeons may be overly cautious. Wouldn't it be good to actually go test it? If NASA goes ahead with the L2 exploration gateway, they will be.

On the other hand, NASA seems dedicated to the medical mitigation of long term zero-g exposure, whereas Zubrin remains an advocate of artificial gravity - particularly using a rotating tether. This may end up being a point he backs down on, especially when NASA starts saying they have the drugs to do a Mars mission.

[Dalhousie]

Tissue cultures can be flown on the ISS for years and should provide much of the data needed on the effects of GCRs.  I am surprised this has not been done as yet.

I agree that the zero gravity issue is well on the way to being solved with a combnation of exrcise, diet, and medication.

[ciscosdad]

You are correct about Zubrin, but I think he may be a little over optimistic in some cases. NASA on the other hand is (perhaps understandably) over cautious.
Zero G seems less of a problem than advertised given that multiple cosmonauts have spent more than 9 months in it with no long term ill effects provided simple precautions are followed.
Given the time of one leg of the Mars expedition is about that and the middle interval is ~500 days on the surface in .38g, there may actually be no problem at all.
Of course if there was an abort midway that forced them to do the return leg immediately, that may be a problem, as I don't believe anyone has spent that amount of time (18 consecutive months) in zero G have they?
I'd like to see a low g (Lunar or Martian) gravity module on ISS. Pity its never been funded. Some tether work in orbit would be useful too.

[Sparky]

Tissue cultures can be flown on the ISS for years and should provide much of the data needed on the effects of GCRs.  I am surprised this has not been done as yet.

I agree that the zero gravity issue is well on the way to being solved with a combnation of exrcise, diet, and medication.

We send large amounts of tissue cultures to ISS all the time. We call them astronauts. But ISS is situated well within the Earth's Magnetic field, which protects from much of the radiation. The radiation environment in LEO is very different than what would be experienced on a deep space mission.

[MATTBLAK]

Artificial gravity with tethers is one of those great ideas in principle. But actually doing this sort of engineering without first practical testing to exhaustive levels will be difficult, not hand-wavingly simple. And VASIMR - a damn cool idea hampered by a couple of sobering facts: How will funding and approval get passed to build a 100 megawatt-plus space worthy nuclear reactor then go fly it?

Zubrin knows this, that's why he has advocated chemically propelled missions. And ISRU? Damned nice to have - maybe even a mission success deal-breaker. But NOT essential in terms of actually going there. Though without ISRU, shortcuts in crew size and mission duration/capability would have to be made, reducing the value of even doing it in the first place. And as for L-2 Gateway Station? I'd say it was almost essential.

[QuantumG]

And VASIMIR - a damn cool idea

What's cool about it? It's just a magnetic mirror with radio frequency heating, pre-tokamak fusion technology that never delivered. Just getting it working for full length burn times requires superconducting magnets which don't exist. Testing it requires vacuum chambers larger than any that exist, or in-space testing. And in the end, you get a thruster that's not much better than existing - flight proven - electric thrusters.

Zubrin knows this, that's why he has advocated chemically propelled missions.

He's advocating chemically propelled missions because he thinks they are the best way to do it. Even if all the ponies fell from the sky to make VASIMR work he'd be advocating chemical. Same with artificial gravity, ISRU and heavy lift.

[MikeAtkinson]

The cancer situation is much more complex than just increased rates due to radiation.

Cancers are caused by multiple factors including lifestyle, diet, exercise, viruses and pollutants and the chances of dying of those cancers depends on age, fitness, cancer type, immune system, early detection and medical care among others. The chances of dying from a cancer changes over time as new treatments are found and risk factors change.

Being accepted onto an astronaut program alters all those risk factors, going on a Mars mission alters them again.

It is by no means certain that radiation ill cause the greatest increase in cancer risk as compared to other factors. And other factors may actually cause a decrease in overall cancer risk. Going on a Mars mission and the greater health monitoring before during and after the mission means that it is highly likely that cancers are picked up earlier than otherwise. Astronauts are likely to receive healthcare second to none for the rest of their lives, meaning that treatments are more likely to be successful than for the general population.

Cancer is complex. Zubrin's naive approach to assessing its risk from Mars missions should not be followed.

[MikeAtkinson]

It is really difficult to predict technology over the long term, so I would not rule VASIMIR out completely.

However from what we know now, VASIMIR seems very unlikely to be useful for early Mars missions. It is certainly not necessary.

This makes it a sideshow. Lets continue to develop the technology together with that for other electric thruster types, but don't put it anywhere on plans and certainly not on the critical path.

One problem with all electric thrusters which hasn't been mentioned so far is the long operational life required, this requires long development, test and qualification runs which makes development slow.

[MATTBLAK]

And VASIMIR - a damn cool idea

What's cool about it? It's just a magnetic mirror with radio frequency heating, pre-tokamak fusion technology that never delivered. Just getting it working for full length burn times requires superconducting magnets which don't exist. Testing it requires vacuum chambers larger than any that exist, or in-space testing. And in the end, you get a thruster that's not much better than existing - flight proven - electric thrusters.

Zubrin knows this, that's why he has advocated chemically propelled missions.

He's advocating chemically propelled missions because he thinks they are the best way to do it. Even if all the ponies fell from the sky to make VASIMR work he'd be advocating chemical. Same with artificial gravity, ISRU and heavy lift.

What's cool about it? I'm surprised I have to point it out to you: the ability to - theoretically - take a crew to Mars in about 40 days. But like you, here's where I would disconnect from the fantasy and join reality. Such huge nuclear reactors for space wont get approved or funded, probably in our lifetimes. Anti-nuke sentiment is far too powerful, therefore, VASIMR and - sadly - nuclear thermal rocket motors wont get approved anytime soon either. Too many people (space cadets?) don't seem to get this fact. And one of the main reasons for Zubrin advocating chemical propulsion is cost - its cheaper and goes hand-in-glove with ISRU.

[Dalhousie]

I agree that the zero gravity issue is well on the way to being solved with a combnation of exrcise, diet, and medication.

We send large amounts of tissue cultures to ISS all the time. We call them astronauts. But ISS is situated well within the Earth's Magnetic field, which protects from much of the radiation. The radiation environment in LEO is very different than what would be experienced on a deep space mission.
[/quote]

The magnetic field protect the astronauts to some extent from the most intense SPEs, they have little or no impact on the GCRs which are far more energenic.  Even the protection from SPEs is only partial, as the particles are trapped to form the VABs.    Since the ISS traverses part of the VABs (the SAA) the average dose is still significant.  The nadir shielding from the Earth is probably more important for the GCR, but it just mans dose times have to be doubled.

Tissue cultures are a way of investigating long term low level dosed well beyond what is practical for crewed mission and complements work from short term high level does in the lab.

However quite a few people have accummulated career doses similar to what you would get on a Mars mission

[MATTBLAK]

Cosmonauts like Sergei Krikalev? I'd love to know what his career dosage has been!

[Dalhousie]

As far as I can determine, annual GCR dose in interplanetary space is ~250 mSv, therefore ~125 mSv in LEO or the surface of the Moon of Mars.

A conjunction-class mission to Mars would expose an astronaut to ~430 mSv.  Krikalev, with 803 days in LEO, has been exposed to 260 MsV.

Because of repeated passes through the SAA he would have accumulated a significant dose from high energy protons and electons - equivalent to higher on average than interplanteray space.  It is hard to get accurate figures, but annual total dose equivalents on the ISS are said to be as low as 180 mSv per year and as high as 400 mSv.

So Krivalev could have accumulated as low as 400 mSv or as high as 880 mSv.

Cumulative dose equivalent estimates on a conjunction mission range between 350 and 950 mSv.

[spectre9]

BAD MDOT IS BAD.

That's all you really need to know about electric propulsion in regards to manned missions.

It's good for super small space probes and super light satellites but not for massive human rated interplanetary spacecraft.

By the time you're jacking out enough thrust to get humans on a high energy 6 month trajectory to Mars the electric drive system be it solar or nuclear is completely unrealistic.

NASA has gone from nuclear thermal to solar electric and it's all just bogus. Nuclear thermal would've been ok if they matured the technology but at this point it's just best to go with storable fuels like those that were used to safely return man from the moon. Important demonstration of hypergols I reckon and should be built on.

[Robert Thompson]

Same with artificial gravity

Is Zubrin dismissing tether architectures?
Favoring accelerate till midpoint, reverse, decelerate till destination?

[MATTBLAK]

To get the best of 'what's around' for Mars:

Earth Departure: LOX/LH2 and/or LOX/Kerosene. Or LOX/CH4 if the buildup of mission hardware is going to take awhile but you still want next-best performance after LOX/LH2. LOX/CH4 is better Isp than LOX/Kerosene, but Hypergols are not much worse. LOX/Ethanol anyone? Technology Risk: Low-to-moderate.

Mars Arrival: Option 1 - Aerocapture then powered & parachute/powered descent. Propellant? Storables, because of the long voyage through interplanetary space: devoting mass and technology to zero-boiloff will reduce Hab/Lander down mass. Option 2 - Direct descent with no aerocapture; parachute & powered descent. Propellant; again storables. Technology Risk: Medium-High.

Mars Ascent to Orbit: Option 1 - Storables all the way, but mission down mass reduced. However, low technology risk. Option 2 - ISRU for LOX by 'cracking' the Martian CO2 atmosphere: solar & RTG-powered ISRU plant. Fuel? Kerosene or Ethanol. Technology Risk: Moderate.

Earth Return Vehicle from Martian orbit: Storables. Technology Risk: Low. This vehicle would be waiting in Martian orbit to do its job for at least two years, so proven hypergolic technology it should be.

*Bonus Category: Launchers - OPTION 1: Mixed fleet of Delta IV-Heavy (uprated to 40+plus tons = known, simple upgrade options), Atlas V, Falcon Heavy and Ariane V. OPTION 2: SLS & Falcon Heavy. Earth Re-entry spacecraft? Either Orion or Dragon, I have no preference.

[MikeAtkinson]

MATTBLAK, I'm not sure what you mean by Technology Risk here, within each of those paragraphs there are several technologies mentioned.

It is better to discuss this in terms of technology readiness level (NASA version)

Aerocapture is probably only TRL 6/7 because although it has been demonstrated in space, those missions were much smaller in terms of mass than a Mars mission would be.

Mars Entry Descent and Landing EDL is difficult to quantify in terms of TRL as demonstrated methods do not scale up the manned mission masses. Red Dragon type approaches could probably scale up to the required size and maybe, just maybe, we might have it demonstrated this decade, then it would be at TRL 7. The DRM 5 method would be sufficient as well, component technologies though are still quite immature.

Technology risk is not the only risk, there are program, financial, safety, cancellation, political (and probably several other risks as well) risks. These need to be balanced in the overall program to get the highest return for the least cost. President+Congress+NASA seem to have been pretty poor at creating manned spaceflight programs which balance all the risks to arrive at something that can be implemented successfully.

[MATTBLAK]

Thanks for your post! I'm aware that Manned Mars Missions themselves are chock-full of Technology Risks - nature of the beast. But my context was for the getting there-and-back propulsion. We can't do or are not ready for nuclear propulsion of any sort. Neither are we ready for Zero-boiloff LOX/LH2 propulsion on the grand scale readiness needed for Mars. And we're not quite there with LOX/CH4 yet.

But we can do storable propellant propulsion on a big scale already, and probably LOX/Kerosene, too.

[laszlo]

So who's Zubrin?

Serious question, not a troll.

Laszlo

[MATTBLAK]

So who's Zubrin?

Serious question, not a troll.

Laszlo

http://en.wikipedia.org/wiki/Robert_Zubrin

http://www.marssociety.org/

[Jim]

We have hunches, and assumptions, but no hard data. Human factors experts and flight surgeons may be overly cautious. Wouldn't it be good to actually go test it?

Mars Odyssey, MRO, and MSL have radiation sensors.

[mmeijeri]

Earth Departure: LOX/LH2 and/or LOX/Kerosene.

I'd say LOX/LH2 for transport of individual modules from LEO to L1/L2 and hypergolics from L1/L2 to TMI. Both segments are part of Earth departure.

Mars Arrival: Option 1 - Aerocapture then powered & parachute/powered descent. Propellant? Storables, because of the long voyage through interplanetary space: devoting mass and technology to zero-boiloff will reduce Hab/Lander down mass. Option 2 - Direct descent with no aerocapture; parachute & powered descent. Propellant; again storables. Technology Risk: Low.

I'd say propulsive insertion into high Mars orbit, possibly a Mars Lagrange point. Much lower technology risk, but also more propellant-intensive. Of course, I like propellant-intensive solutions. The reason for braking into high rather than low Mars orbit is that it requires much less propellant. We also need to avoid outgrowing a transfer stage that will still fit inside an EELV fairing.

Mars Ascent to Orbit: Option 1 - Storables all the way, but mission down mass reduced. However, low technology risk. Option 2 - ISRU for LOX: solar & RTG-powered ISRU plant. Fuel? Kerosene or Ethanol. Technology Risk: Moderate.

I doubt avoiding ISRU (and surface nuclear power) is practical, and using LOX/hydrocarbon should be easier than hypergolics.

Earth Return Vehicle from Martian orbit: Storables. Technology Risk: Low. This vehicle would be waiting in Martian orbit to do its job for at least two years, so proven hypergolic technology it should be.

Agreed.

*Bonus Category: Launchers - OPTION 1: Mixed fleet of Delta IV-Heavy (uprated to 40+plus tons = known, simple upgrade options), Atlas V, Falcon Heavy and Ariane V. OPTION 2: SLS & Falcon Heavy. Earth Re-entry spacecraft? Either Orion or Dragon, I have no preference.

The main reason I see for an EELV Phase 1 is ACES the EDS, with the bigger launcher as an inconsequential side-effect. No need to avoid use of commercially available large launchers, but no real need for such use, and not much benefit either.

[mmeijeri]

Mars Odyssey, MRO, and MSL have radiation sensors.

Important data, but it doesn't assess the effect of various designs of radiation shielding or the biological end effects of the internal radiation environment. Tissue cultures could help with that.

[Jim]

Mars Odyssey, MRO, and MSL have radiation sensors.

Important data, but it doesn't assess the effect of various designs of radiation shielding or the biological end effects of the internal radiation environment. Tissue cultures could help with that.

how do you know about the shielding aspects?

[mmeijeri]

how do you know about the shielding aspects?

I meant testing various designs of shielding for manned spacecraft to reduce the biological end effects. Were you suggesting there are (physical) radiation shielding experiments aboard the probes you mentioned, rather than just measurements of the local radiation environment?

[mmeijeri]

And as for L-2 Gateway Station? I'd say it was almost essential.

An MTV parked at L1/L2 would essentially be a gateway station.

[QuantumG]

What's cool about it? I'm surprised I have to point it out to you: the ability to - theoretically - take a crew to Mars in about 40 days.

It's not free - that's Zubrin's point. Even if you could take a crew to Mars in about 40 days, you have to pay for it with decreased crew safety. If your reason for wanting to go faster is better crew safety then you've just negated your goal.

(I've explained it three times now).

[go4mars]

Is Zubrin dismissing tether architectures?

No.  Zubrin is a proponent of tether architectures so that people arrive strong. 

He's built some pretty compelling ISRU demonstrators. 

Back to the original post, my belief is that his most pregnant ideas yet to be adopted in the main stream as obviously good ideas are "transorbital railroad" and "nuclear salt water".

[douglas100]

Zubrin has some interesting ideas and is very good at thinking outside the box. But...

Nuclear salt water has never been done. It would have to be developed and tested. Controlling the fission reaction in such a dynamic system would be far from simple. The exhaust products would be so radioactive as to make testing on Earth very difficult and expensive. And politically it's difficult to imagine this being acceptable. Even getting the propellant into space where the system would have to be be used, would be hazardous: imagine what would happen during a launch accident where the payload is many tons of uranium salt solution. I don't see being developed.

The transorbital railroad proposal is essentially a government launch subsidy. I've no idea what the political feasibility of such a scheme would be. Whether it would stimulate the launch market remains to be seen. What effect it might have on international trade agreements is another open question.

I think you're right in saying that Zubrin's ISRU ideas are his best. Some are almost mainstream already.

[pathfinder_01]

Ah the lone voice in the wilderness here. 

Electric propulsion is fine, even solar electric is fine for cargo or delivering propellant to mars.

It just gets iffy when you want it to move the crew to Mars with it as main propulsion. Then you may need nuclear power.

You could also use it to move a chemically powered Mars spacecraft assembled in LEO out to L1/l2(or a high earth orbit) reducing the mass of chemical propellant needed to get to mars and perhaps the overall mass of the mission.  That ability could reduce the launch costs of such a spacecraft.

[A_M_Swallow]

And as for L-2 Gateway Station? I'd say it was almost essential.

An MTV parked at L1/L2 would essentially be a gateway station.

True but we would still want to make trips to the Moon whilst the MTV was at Mars.

I suspect that the initial EML Gateway Station can be smaller than the ISS.

[MATTBLAK]

What's cool about it? I'm surprised I have to point it out to you: the ability to - theoretically - take a crew to Mars in about 40 days.

It's not free - that's Zubrin's point. Even if you could take a crew to Mars in about 40 days, you have to pay for it with decreased crew safety. If your reason for wanting to go faster is better crew safety then you've just negated your goal.

(I've explained it three times now).

With this post, you've selectively ignored what I wrote immediately afterwards! This is not the Monty Python argument sketch you know!!

[mmeijeri]

True but we would still want to make trips to the Moon whilst the MTV was at Mars.

You could keep a spare at L1/L2. Call that a gateway station if you will, but my point was that you don't need to design a separate station or even build a spare if you can't afford it. Similarly, a lander could be its own makeshift gateway station.

[go4mars]

I know Bob prefers nuclear power on the surface of Mars, and chemical for  propulsion...   Are there major show-stoppers for designing a fission power generator that would be pretty good for both?

[go4mars]

Asking me?  No I don't mean a Mars transfer vehicle / cycler.  I mean a reactor you would use for inter-space propulsion on the way there, and also on the surface of Mars.  Same unit.  It's use on the way there would probably need to be merely an auxillary to the chemical prolusion.

It's been a number of years since I've read "The Case for Mars".  But I know he didn't want to waste time developing an in-space advanced propulsion system.  But he wanted nuclear for the surface.  Presumably, he considered the power to weight ratio as not worthwhile or maybe had safety reservations. 

Now that I'm thinking about it: using his Mars Direct architecture, if it arrives on the preceding conjunction with the earth return vehicle anyways, and the manned lander is the heavier chunk, then I guess it becomes more easy to see that the incorporating it to help on the journey doesn't make a lot of sense.  Added comlexity and risk. 

Edit add: just noticed to context I hadn't paid attention to up the thread.
Was assuming power use could be made to be similar.  Will need to consider further.  And revisit that book of his.

[savuporo]

MGB should have been flown and supported much better, and something along its lines should be done for radiation as well, outside the van Allen belts of course. Maybe now that we actually have a working microfinancing platform, it should be tried again.
All these discussions around mars architectures would be far less theoretical if such basic ( and incredibly cheap ) data acquisition missions were flown.

[MATTBLAK]

True but we would still want to make trips to the Moon whilst the MTV was at Mars.

You could keep a spare at L1/L2. Call that a gateway station if you will, but my point was that you don't need to design a separate station or even build a spare if you can't afford it. Similarly, a lander could be its own makeshift gateway station.

Re-usable MTV? With storables, it takes the crew to Mars for either propulsive braking into High Mars orbit (DEIMOS!!) or aerocapture. Then it waits for the crew to return and head back to Earth for propulsive braking into L-2. To do all that would take a heck of a lot of propellant, but at least between TMI windows there would be lots of time to fill it back up with a 'Propellant Railroad' from Earth. Maybe its propulsion/propellant module could be detachable; dock a new one to it between each Mars mission or when its design life is expired, whichever comes first. Engines? I suggest a quartet of AJ-110 derived motors.

How big would an empty tank/engine module for a mission like this be if you had to launch it on an existing EELV? Or would it have to be launched in two separate engine & tank packages?

[MATTBLAK]

Asking me?  No I don't mean a Mars transfer vehicle / cycler.
Edit add: just noticed to context I hadn't paid attention to up the thread.
Was assuming power requirements would be similar.  Will need to consider further.  And revisit that book of his.

Sorry: was replying to mmeijeri! Have changed previous post to clarify that.

[Kaputnik]

And ISRU? Damned nice to have - maybe even a mission success deal-breaker. But NOT essential in terms of actually going there. Though without ISRU, shortcuts in crew size and mission duration/capability would have to be made, reducing the value of even doing it in the first place. And as for L-2 Gateway Station? I'd say it was almost essential.

I would put those the other way around.
Show me how you put together a non-ISRU Mars Ascent Vehicle without needing some Sea-Dragon sized LV to get it there.

My 'wish list' for technology developments the enable a Mars mission goes:
ISRU
Large scale EDLS
Surface power, EVA ops, and hab systems
In-space hab systems

Those are the real breakthrough developments that we need. Everything else, from aerocapture to L2 stations to super-LVs, are just 'nice to haves'.

[MATTBLAK]

"Show me how you put together a non-ISRU Mars Ascent Vehicle without needing some Sea-Dragon sized LV to get it there."

Well, you are overstating things a tad:a non-ISRU Ascent Vehicle need not be a 'Battlestar Galactica' sent there on a rocket twice as powerful as Ares V!! First of all, I only said that a non-ISRU Ascent Vehicle was an option, and not the preferred one at that.

Following figures approximate; since Mars mission mass numbers are still debated - The dry mass of the Mars Ascent Vehicle (MAV)  according to the NASA Mars DRM-3 study - is 20 tons, give or take consisting of both the Descent and Ascent portions, plus 20 tons of cargo sent as down mass:

http://ares.jsc.nasa.gov/HumanExplore/Exploration/EXLibrary/docs/MarsRef/addendum/index.htm

This down mass includes the ISRU 'feedstock', a Rover, scientific equipment, tools, spares, the nuclear reactor for ISRU etcetera. If this vehicle is landed with 6 tons of LH2 feedstock for ISRU creation of LOX/Methane, this is leveraged to produce more than 30 tons of propellants to launch the ascent vehicle to a Mars Transfer Vehicle or a dedicated Earth Return Vehicle in Martian orbit. We're assuming that the MAV would use this approx 30 tons of propellant to liftoff only the crew cabin, the Ascent stage motor, the crew and their Martian regolith & rocks, all massing about 15 tonnes at Martian liftoff - leaving the legs, descent motors and cargo pallet behind.

Since hypergolic propellants are between 10 and 15% percent less efficient than LOX/CH4, if we halved the vehicle's cargo downmass load - including deleting the nuclear reactor - we could more than make up for the loss in efficiency.

But if the Ascent Vehicle is made as basic and bare-bones as possible: a small cabin derived from Dragon or another, all-composite capsule, bolted to an Ascent engine, RCS sets and fuel tanks. Think of it as a "Apollo Lunar Module Ascent Stage on Steroids".

Apart from the hypergolic propellant load, all this craft would be lifting to Martian orbit would be a crew (2 or 3?) and their load of Martian rocks and regolith and maybe some data storage blocks. So it doesn't have to be as big or complex as the 44 ton (landed), 15 ton (ascending) NASA DRM-3 design for 6 crew.

Or if you wanted to meet ISRU halfway and only produce LOX oxidizer to supplement some landed kerosene or ethanol - this could use a 5 metric ton combined solar and RTG package to slowly produce the LOX only. And you could trade some of that 5 tonne increase in mass for cargo down mass; as you'd be using more powerful LOX/Kerosene/Ethanol for descent propulsion, not less efficient storables, to compensate for the mass of the ISRU package in the first place.

But let me state for the record: I am in favour of ISRU for Mars.

But an all-hypergolic Ascent Vehicle doesn't have to weigh 60 or 70 tonnes at Trans-Mars Injection. Launched on an uprated EELV or Falcon Heavy, the Hypergolic 'basic' MAV with Empty propellant tanks would mass about 16 metric tons, Descent/Ascent stages all up with little or no cargo. I say make a separate, dedicated Cargo Lander by deleting the weight of the 4 ton composite crew cabin. Before Earth Departure, load it with fuel from a Prop. Depot or Propellant Delivery Modules until it contains enough hypergolics to descend to the Martian Surface. For a bare-bones 2 or 3 person Ascent vehicle weighing half that of the 6 person DRM-3 design, total hypergolic propellant loading would be about 35 tons - 15 tons for a descent assisted by frictional slowing and parachutes, plus 20 tons to later get the craft into orbit to meet the ERV or MTV.

As I said, by making the MAV basic with no 'bells & whistles' and keeping bulk cargo for another lander, you don't need a 'Sea Dragon' to send it to Mars!! A couple Phase 1 EELVs, Falcon Heavies or one SLS Block 1B could do it. But yes, if you wanted a 6x person version with lots of cargo down mass, expect a much heavier vehicle. If you wanted to lift a hypergolically propelled big, 6 person MAV then yes, get a bigger rocket than even Ares V would have been.

[A_M_Swallow]

{snip}
But if the Ascent Vehicle is made as basic and bare-bones as possible: a small cabin derived from Dragon or another, all-composite capsule, bolted to an Ascent engine, RCS sets and fuel tanks. Think of it as a "Apollo Lunar Module Ascent Stage on Steroids".

The SEV weights about 3 tonnes and supports 2 men.  It can be used as the lander's cabin, to move around on Mars and the Ascent stage's cabin.
http://www.nasa.gov/pdf/464826main_SEV_FactSheet_508.pdf

If your stores are landed separately the Space Exploration Vehicle (SEV) is the idea vehicle to go and get them.  Possible using the MMSEV version.

Apart from the hypergolic propellant load, all this craft would be lifting to Martian orbit would be a crew (2 or 3?) and their load of Martian rocks and regolith and maybe some data storage blocks. So it doesn't have to be as big or complex as the 44 ton (landed), 15 ton (ascending) NASA DRM-3 design for 6 crew.

Or if you wanted to meet ISRU halfway and only produce LOX oxidizer to supplement some landed kerosene or ethanol - this could use a 5 metric ton combined solar and RTG package to slowly produce the LOX only. And you could trade some of that 5 tonne increase in mass for cargo down mass; as you'd be using more powerful LOX/Kerosene/Ethanol for descent propulsion, not less efficient storables, to compensate for the mass of the ISRU package in the first place.
{snip}

The waste product from electrolysing CO₂ is carbon monoxide, a medium power fuel in its own right.  Lowish Isp means that bigger but empty pair of fuel tanks are needed.

[A_M_Swallow]

True but we would still want to make trips to the Moon whilst the MTV was at Mars.

You could keep a spare at L1/L2. Call that a gateway station if you will, but my point was that you don't need to design a separate station or even build a spare if you can't afford it. Similarly, a lander could be its own makeshift gateway station.

Re-usable MTV? With storables, it takes the crew to Mars for either propulsive braking into High Mars orbit (DEIMOS!!) or aerocapture. Then it waits for the crew to return and head back to Earth for propulsive braking into L-2. To do all that would take a heck of a lot of propellant, but at least between TMI windows there would be lots of time to fill it back up with a 'Propellant Railroad' from Earth. Maybe its propulsion/propellant module could be detachable; dock a new one to it between each Mars mission or when its design life is expired, whichever comes first. Engines? I suggest a quartet of AJ-110 derived motors.

The station keeping thrusters and fuel tanks on LEO and EML spacestations probably do not need to be as big as the main engine on a MTV.

However commonality between the cabins, docking systems and life support sounds like a good idea.

How big would an empty tank/engine module for a mission like this be if you had to launch it on an existing EELV? Or would it have to be launched in two separate engine & tank packages?

Since the spacecraft is in a vacuum it does not need to be thin.  Attaching drop tanks to the side provide a simple way of both fuelling the vehicle and reducing the return mass.

[Kaputnik]

Mattblack- thanks for the reply. I think we are coming at this from quite different POV on what is an achievable landed mass.
Given that current state of the art is about 1t landed, I tend to envisage anything up to 10t as a reasonably scaling up of heritage technology, with 20t as a likely upper limit before fundamentally different approaches must be taken.
A bare-bones ISRU ascent vehicle could, IMHO, be squeezed into such a mass (assuming multiple landers to fly the crew, the ISRU plant, etc). A pre-fuelled ascent vehicle clearly could not.

I should also add that the multiple small landers approach has various benefits. All of the more advanced EDLS technologies are likely to lead to a reduced landed mass fraction; therefore the IMLEO of the mission is likely to be higher. That's not to mention the costs and engineering challenges inherent in such development.

[mmeijeri]

I say make a separate, dedicated Cargo Lander by deleting the weight of the 4 ton composite crew cabin. Before Earth Departure, load it with fuel from a Prop. Depot or Propellant Delivery Modules until it contains enough hypergolics to descend to the Martian Surface.

Why before TMI (even from L1/L2 rather than LEO) instead of in high Mars orbit? You could then send the propellant by SEP and maybe be more mass-efficient than all LOX/LH2.

[MATTBLAK]

Yes! Good point.

[Robotbeat]

I say make a separate, dedicated Cargo Lander by deleting the weight of the 4 ton composite crew cabin. Before Earth Departure, load it with fuel from a Prop. Depot or Propellant Delivery Modules until it contains enough hypergolics to descend to the Martian Surface.

Why before TMI (even from L1/L2 rather than LEO) instead of in high Mars orbit? You could then send the propellant by SEP and maybe be more mass-efficient than all LOX/LH2.

Even better is to send the full, fueled lander to low Mars orbit (or high Mars orbit, whichever is where you plan on the crew getting on board) via SEP tug.

[mmeijeri]

Even better is to send the full, fueled lander to low Mars orbit (or high Mars orbit, whichever is where you plan on the crew getting on board) via SEP tug.

Maybe if you have a SEP tug that can transport a full lander, but an empty lander has lower technology requirements. The reason I was looking at high Mars orbit is because I imagined that's where the crew would leave the MTV and enter the lander. Of course that doesn't preclude additional refueling of the lander in low Mars orbit.

[MATTBLAK]

Maybe Deimos could be the staging area for arriving Earth  Return Vehicles and Ascent vehicles, sent there by SEP tugs first.

[DLR]

What's the advantage of Deimos orbit over any other Mars orbit?

[MATTBLAK]

What's the advantage of Deimos orbit over any other Mars orbit?

Low delta-vee (propellant) requirements to reach it - as you approach Mars, you would not yet be close enough to be greatly accelerated by the Martian gravity, requiring lots of fuel to thrust into Mars orbit. Also, Deimos is in direct line of sight communications with Earth most of the time and each hemisphere of Mars for a goodly portion of each Sol (Martian day). Deimos would be a good place to Tele-robotically operate Rovers, landers and sample return probes.

[Hop_David]

What's the advantage of Deimos orbit over any other Mars orbit?

Low delta-vee (propellant) requirements to reach it - as you approach Mars, you would not yet be close enough to be greatly accelerated by the Martian gravity, requiring lots of fuel to thrust into Mars orbit or heavily aerocapture. Also, Deimos is in direct line of sight communications with Earth most of the time and each hemisphere of Mars for a goodly portion of each Sol (Martian day). Deimos would be a good place to Tele-robotically operate Rovers, landers and sample return probes.

.7 km suffices for a Mars capture. If Mars periapsis is in the upper atmosphere, aerobraking will shed velocity each orbit, lowering apoapsis.

So with the use of aerobraking, a low Mars orbit can be achieved with .7 km/s. A Mars ascent vehicle could rendezvous with something in LMO much easier than a Deimos orbit.

Not including gravity loss, it'd take about 5.7 km/s for a MAV to rendezvous with Deimos. It'd take about 3.7 km/s for an MAV to rendezvous with something in LMO.

[MATTBLAK]

Sorry - my bad: I was supposed to be comparing Deimos with Phobos delta-v in terms of ease of reach. Aerobraking will always use less propellant for capture than propulsive! (previous post edited).

[MATTBLAK]

Hah! In the various posts I've been doing around Mars-based threads around here, I've often mentioned a more basic manned Mars mission with RTG & Stirling based power plants, smaller crews and mixed storable/LOX-methane propulsion - only to find somebody already did a very fine paper on the subject. The only major difference is; when this 'Austere Mars Mission' paper was authored, Ares V was still in the picture. I've been advocating uprated Delta IV-H, Atlas V & Falcon Heavy mixed launch fleets for the mission, with SLS as a 'nice-to-have-if-funding-is-there' backup option.

http://trs-new.jpl.nasa.gov/dspace/bitstream/2014/41431/1/09-3642.pdf

[Hop_David]

What's cool about it? I'm surprised I have to point it out to you: the ability to - theoretically - take a crew to Mars in about 40 days.

40 day VASIMR trip relies on an implausibly high specific power.

[MATTBLAK]

You are quite right. Er... But I'm a bit puzzled: every time I have mentioned VASIMR in this thread, a couple people seem to think I was advocating and defending it strongly! You can see from the quoted post that I bolded & italicized theoretically. And I've told Quantum more than once that I doubted a big enough nuclear reactor to power an effective VASIMR engine would ever be approved. On that basis, I have dismissed it as a contender, even in the 'nice to have' ranks.

[QuantumG]

Why can't some people spell VASIMR? I've noticed various misspellings everywhere...

[MATTBLAK]

You're quite right! I keep thinking of how I've heard it pronounced out loud so many times: "VAS-I-MEER".

I'll try not to keep subconciously typing it that way.

EDIT: There! I've gone back through this thread and changed the spelling of the word each time I used it. But I don't have time to go and change every time I typed it on this whole website: Even I'm not that pedantic!

[Patchouli]

What's cool about it? I'm surprised I have to point it out to you: the ability to - theoretically - take a crew to Mars in about 40 days.

40 day VASIMR trip relies on an implausibly high specific power.

Even with more reasonable  6 to 12MW power which should be achievable even with solar it would revolutionize the mission design.
http://www.adastrarocket.com/Andrew-SPESIF-2011.pdf
6 MW gets the round trip down to 13 months which is within the limits of what is known as far as mission duration goes.
Mass optimized Ultraflex arrays can make up to 500W per kg thin film can do even better.
http://www.sbir.gov/sbirsearch/detail/16067
Just ultraflex would put you way over the power to mass needed as the 6 MW array even accounting for Mars distance would only be a measly 33 metric tons.
Even if I double this to 66 tons to account for extra support structure power conditioning and redundancy the mass numbers still look good compared to a conventional mission.

[QuantumG]

6 MW gets the round trip down to 13 months which is within the limits of what is known as far as mission duration goes.

Do you have any idea what the largest space nuclear power source ever flown is?

[Patchouli]

6 MW gets the round trip down to 13 months which is within the limits of what is known as far as mission duration goes.

Do you have any idea what the largest space nuclear power source ever flown is?

My masses were for solar not nuclear.
BTW I know the largest space reactor flown TOPAZ 2 was only 6 KW but it used  thermal ionic conversion which was about 5.2% efficient.
A Rankine cycle turbine can easily achieve 40+% efficiency.

So if you swapped in a turbine but kept the reactor the same size you could boost the power output by a factor of up to eight bringing the power output to 48.46 KW.

Even here the reactor mass should be 143,184Kg.
Solar totally wins here but the mass numbers are good enough for Zurbin's ISRU fuel plant which only needs 80 to 100 KW as only two 1061Kg reactors would be needed this number includes the cooling system.

it should be noted they didn't really push for power to weight with TOPAZ instead reliability and simplicity were the primary goals.
A molten salt reactor would get much better numbers.

[QuantumG]

6 MW gets the round trip down to 13 months which is within the limits of what is known as far as mission duration goes.

Do you have any idea what the largest space nuclear power source ever flown is?

My masses were for solar not nuclear.

Fine.. do you have any idea what the biggest solar power source ever flown is? (and the power to mass of that system?)

If you believe in solar powered plasma rockets, you might as well believe in solar power satellites too. Zubrin calls it "an alternate universe".

[DLR]

What has been flown in the past is not important.

Solar cell technology has advanced quite considerably even since the ISS's wings were put up. Solar tugs will be a viable space transportation option.

And you can't seriously compare Topaz to more capable space nuclear power reactor designs.

[QuantumG]

Until it has flown it's just powerpoint.

[savuporo]

Fine.. do you have any idea what the biggest solar power source ever flown is? (and the power to mass of that system?)

Its by far not the biggest, but Ultraflex thin film arrays mentioned above flew on Phoenix

"The Phoenix Ultraflex arrays are the highest performance space-rated solar arrays ever used. The specific power for these arrays is greater than 105 W/kg,

[QuantumG]

Its by far not the biggest, but Ultraflex thin film arrays mentioned above flew on Phoenix

"The Phoenix Ultraflex arrays are the highest performance space-rated solar arrays ever used. The specific power for these arrays is greater than 105 W/kg,

Cool, so you only need 57 tons of solar panels to do this 6MW mission.

[savuporo]

Cool, so you only need 57 tons of solar panels to do this 6MW mission.

No, because array specific power scales well with size ( square law and all that ), and Phoenix was quite small.
Plus, efficiencies are ever slightly increasing, as demoed on FTSCE

EDIT: Apparently UltraFlex config has hit 175w/kg in larger version already, with presumably even higher design goal for Orion.

EDIT2: and of course then there are all the SBIR's listed by Robotbeat here that may or may not hit anywhere from 200w/kg to 500w/kg at system level at some point.

[Dalhousie]

Sprint missions (~12 months round trip) are always interesting because of their short mission times, and it is good to see that latest solar technology linked to VASIMR is starting to make them look half way feasible.

The trouble is they still require 300-400 tonnes in LEO and their return - one month of the surface - is low compared to a chemicalfueled conjunction mission which would spend 18 months on the surface for 2-3 times the mass and less development.

The whole point of going to Mars is what you can do on the surface, propulsion is simply a means to that end, not an end in itself.

[Warren Platts]

NASA adopts the best of Zubrin's ideas, eventually

Once upon a time, it seemed that Zubrin's name was unspeakable at NASA.. and you can understand why.

They used to dismiss the leverage of Mars ISRU, but now it's a standard part of any architecture.

The problem is his best Mars ideas aren't that great. He is the guy that gave us Ares V, after all. His fundamental "insight" is that it is possible to reject Greason's planet hopping strategy and still get to Mars if everything is staged out of LEO. The result, however, is an expensive, mass-starved, shoestring, non-reusable architecture that requires all sorts of engineering fixes. While Zubrin is a nice guy and a great speaker who I always enjoying listening too (and I am very intrigued by his ideas on developing methanol ISRU for use on Earth), arguably the Mars Direct idea and it's numerous offspring have done more than any other single idea to set back NASA's HSF program.

[HappyMartian]

...
While Zubrin is a nice guy and a great speaker who I always enjoying listening too (and I am very intrigued by his ideas on developing methanol ISRU for use on Earth), arguably the Mars Direct idea and it's numerous offspring have done more than any other single idea to set back NASA's HSF program..

Well, Zubrin had lots of help...


From: Wesley T. Huntress: Author of NASA's New Strategy
Questions for the originator of the "stepping-stones to Mars" approach
By James Oberg   April 2010
At: http://spectrum.ieee.org/aerospace/space-flight/wesley-t-huntress-author-of-nasas-new-strategy
"The agency is to send people to Mars using a series of ‘stepping-stone’ destinations that are themselves of interest: Lagrange points, near-Earth asteroids, and Martian moons. The plan is pretty much exactly what Planetary Society president Wesley T. Huntress Jr. proposed in 2004."


And:


From: Enabling Profitable Asteroid Mining
At: http://www.nasa.gov/pdf/383154main_53%20-%2020090803.7.toAugustineCommittee-2009-08-03.pdf
Submitted to Mr. Norman R. Augustine, Chair Review of United States Human Space Flight Plans Committee
August 3, 2009
By Abundant Planet
Board of Directors
William BC Crandall, MBA
Founder, Abundant Planet
Larry Gorman, Ph.D.
Professor of Finance
Cal Poly, San Luis Obispo
Peter Howard, Ph.D.
Senior Scientist, Exelixis, Inc.
South San Francisco, CA
Board of Advisors
Frans von der Dunk, Ph.D
Professor of Space Law
University of Nebraska-Lincoln
Martin Elvis, Ph.D
Senior Astrophysicist, Harvard-
Smithsonian Center for Astrophysics
Dante Lauretta, Ph.D.
Director, Southwest Meteorite Center
University of Arizona, Tucson
Jordi Puig-Suari, Ph.D.
Professor of Aerospace Engineering
Cal Poly, San Luis Obispo
Hon. Andrea Seastrand
Executive Director
California Space Authority
Mark Sonter, MAppSc
Asteroid Enterprises, Pty Ltd
Queensland, Australia
“Rather than the resource-poor moon, or too distant Mars, our national civil space agency should focus on asteroid mineral resources—specifically the platinum group metals. The terrestrial sources of these vital metals are in rapid decline.”


How did we space cadets ever let the "Abundant Planet Board of Directors" get by with their false and silly claim of "the resource-poor moon"?


So maybe a few other folks helped "set back NASA's HSF program". Perhaps, if folks wanted to be honest about it, we space cadets also made some contributions to the "set back" by failing to advocate for a doable and logical approach to "NASA's HSF program"... However, the blame game will never get anyone back to to the Moon.


If a space cadet wants affordable and sustainable missions to Mars, Ceres, and NEOs, gaining access to Lunar ice and propellant makes a whole lot of sense. 


From: Moon Express Delivers Lunar Mission Design Report to NASA Detailing technical plans toward mining the Moon for precious planetary resources   April 23, 2012
At: http://www.prnewswire.com/news-releases/moon-express-delivers-lunar-mission-design-report-to-nasa-detailing-technical-plans-toward-mining-the-moon-for-precious-planetary-resources-148538005.html
"'The Moon has never been explored from an entrepreneurial perspective,' said Moon Express co-founder and chairman, Naveen Jain. 'Think of the Moon as the Earth's eighth continent, potentially the largest repository of asteroid resources in the solar system, and we have barely begun to explore it.'"



Cheers!

[anonymous]

Sprint missions (~12 months round trip) are always interesting because of their short mission times, and it is good to see that latest solar technology linked to VASIMR is starting to make them look half way feasible.

The trouble is they still require 300-400 tonnes in LEO and their return - one month of the surface - is low compared to a chemicalfueled conjunction mission which would spend 18 months on the surface for 2-3 times the mass and less development.

The whole point of going to Mars is what you can do on the surface, propulsion is simply a means to that end, not an end in itself.

There's nothing about solar-electric that requires you to only spend a month on Mars. After a few early missions doing that you can switch to conjunction-class missions and spend 18 months on the surface too.

[QuantumG]

There's nothing about solar-electric that requires you to only spend a month on Mars. After a few early missions doing that you can switch to conjunction-class missions and spend 18 months on the surface too.

Wow, did you fail orbital mechanics or did you just not take it?

If you want faster than ~5 month transit times you must fly an opposition class mission. Zubrin's argument is that there's no abort modes for opposition class missions.

[pathfinder_01]

There's nothing about solar-electric that requires you to only spend a month on Mars. After a few early missions doing that you can switch to conjunction-class missions and spend 18 months on the surface too.

Wow, did you fail orbital mechanics or did you just not take it?

If you want faster than ~5 month transit times you must fly an opposition class mission. Zubrin's argument is that there's no abort modes for opposition class missions.

This is one of thoose things where it may not be comparable. Electric propulsion typically has longer launch windows than chemical, not sure if that is in effect or if it is just because you have a "amazing" power source that lets you get to mars quick enough with a manned mission.  Ad astra thinks if you had 12mw and achived 4kg/kw you could do it in three months, but the viability of such a power source using solar panels is in question.

[QuantumG]

I wish people would provide references when attributing something to Ad Astra.. I know they make a lot of absurd claims, but it'd be great to be able to actually quote them.

[Warren Platts]

Ad astra thinks if you had 12mw and achived 4kg/kw you could do it in three months

If you guys seriously gotta propose 12 MW SEP's the size of skyscrapers, that's some serious giggle factor to get past. It appears to be grasping at straws. I mean I don't even get the point. What's so hard about doing a Mars mission with all conventional chemical rockets?

[pathfinder_01]

I wish people would provide references when attributing something to Ad Astra.. I know they make a lot of absurd claims, but it'd be great to be able to actually quote them.

It is for nuclear power(not solar) but it would not matter the power source used(just the mass/KW).

http://www.adastrarocket.com/VASIMR_for_flexible_space_exploration.pdf

http://www.adastrarocket.com/Andrew-SPESIF-2011.pdf

The 200MW , 40 day one is fiction pure fiction with current technology but 10-12MW from Nuclear might be achievable, Solar I am unsure of all the factors involved.

[QuantumG]

No.. the conjunction class fast transit nonsense.. where'd you get that from?

[Dalhousie]

There's nothing about solar-electric that requires you to only spend a month on Mars. After a few early missions doing that you can switch to conjunction-class missions and spend 18 months on the surface too.

Absolutely.  I was refering to the specific VASIMR proposal referenced, of a one year round trip with 12 MW and a one month stay.

[Dalhousie]

Wow, did you fail orbital mechanics or did you just not take it?

If you want faster than ~5 month transit times you must fly an opposition class mission. Zubrin's argument is that there's no abort modes for opposition class missions.

Conjunction and opposition are not the only options. Other possibilities are:

Sprint missions (12 months, short stay)

Low thrust missions (400-600 day short stay).

Also flyby lander missions (700-800 days, short stay).

And of course the cyclers (essentially using multiple flyby spacecraft to enable long stays on Mars).

I don't particlarly like any of these options (nearly all short stay), but they are possible.

[pathfinder_01]

No.. the conjunction class fast transit nonsense.. where'd you get that from?

One of the things I enjoy learning about is electric propusion. One of the problems is that comparing the two Chemical vs. Electrical isn't as simple as you would think.  Basically Electric propusion uses different trajectories than a chemical rocket would use.

That is why I am not too sure you can compare an electric mission to a chemical mission by conjunction class. An electric mission can leave at any almost any time (although doing so may take much longer to get to your destination than leaving at a better time), where as chemical and nuclear usually cannot. The result is longer launch windows on one hand. In addition electric propulsion can achieve much higher delta V than chemical or some forms of nuclear propulsion. 

What bites is when you want to go into orbit, and then the high thrust of chemical or nuclear makes more sense (which is why the Ad Astra mission goes for direct reentry leaving the spacecraft  before it gets in orbit).

[QuantumG]

 

First time I've used that emoticon. Can I suggest that you "learn" more and come back then? This is getting absurd.

[Dalhousie]

This paper by Wooster et al. shows dV requiements for conjunction class missions with transit times between 120 and 270 days.

http://www.marsjournal.org/contents/2007/0002/files/wooster_mars_2007_0002.pdf

[Patchouli]

Cool, so you only need 57 tons of solar panels to do this 6MW mission.

Considering how massive all the propellant a chemical or even a NTR mission needs that is not bad at all.
http://www.nasa.gov/pdf/373665main_NASA-SP-2009-566.pdf
The Mars DRA 5.0 needed 7 Ares Vs and one Ares I vehicle.
Just the hab and it's EDL systems is going to be more massive then 57 tons.
The total mass for the SEP tug would be half that of the NTR stage and probably a third that of a chemical stage of equal payload.

[Warren Platts]

But the chemical stage would be a lot cheaper to develop and manufacture and assemble in orbit. It all depends on the cost of propellant.

[Patchouli]

But the chemical stage would be a lot cheaper to develop and manufacture and assemble in orbit. It all depends on the cost of propellant.

Really depends on the cost of flight rate of the LV.
You still need depot technology to make it work as well.
But for cargo at least it seems SEP wins hands down it can bring two or three times the mass to Mars for a given IMLEO payload as the next best thing.
Of course if you have a cheap LV and depots mass fraction becomes a lot less of an issue.

[QuantumG]

Cool, so you only need 57 tons of solar panels to do this 6MW mission.

Considering how massive all the propellant a chemical or even a NTR mission needs that is not bad at all.

Wow.. do you know what means?

[JohnFornaro]

We send large amounts of tissue ... all the time. We call them astronauts.

Too funny. I have a box of astronauts in the bathroom even as we speak.  Complete with a pretty floral pattern...

[Solman]

 Solar thermal/electric - big concentrator (part of whose whose support structure doubles as radiator) - solar thermal rocket and solar concentrator type PV that can rotate into and out of the focus to power a MPD or other electric propulsion system. Best of both worlds - LEO to escape in a week or two and then electric to reduce transit time and open up more launch windows.

[Solman]

 Just want to add that perhaps lithium or lithium hydride could be used as propellant for the solar thermal rocket instead of hydrogen. No large tank and non cryogenic but still better than chemical. Lithium also for the MPD or other electric propulsion. Alternately, an arc jet could be used to increase Isp with perhaps no MPD. The solar thermal could use 20% or so of the energy at the focus to vaporize and heat the propellant and leave the rest to the PV to power the arc jet to roughly double the 520? sec. of the lithium or lithium hydride to over 1000 sec.
 For ISRU perhaps oxygen since its so abundant in NEO's and the Moon and easy to extract (using the concentrator for a solar furnace) if it can work with an arc jet.   

[Chris Bergin]

Right folks, this gets back on topic as of now.

[baddux]

So back on topic:

I think that the fact that Orion exists and will be there will block any sensible Mars plans for years. They need to use Orion in their plans because they have spent 1.5 billion or something to it and it is very much moon-specific vehicle, no use in Mars missions. No sense having 25mt vehicle for just LAS and re-entry (orion could not even be any sort of LAS vehicle), that could be done in 4mt dragon-type vehicle.

Btw if I were Dr. Zubrin I probably would create a NEO mission plan with mars direct style and promote it. He has already said in some speech that NEO mission is fine because you can do it with subset of Mars mission (probably like MD hab module and SEP in the garage etc.). That would be more achievable than straight to Mars and new president could say that we will go to a NEO in 10 years from now and do the other things...

[JohnFornaro]

So back on topic:

I think that the fact that Orion exists and will be there will block any sensible Mars plans for years. ...

That statement doesn't make a lot of sense to me, since Orion seems to be an integral part of how they plan to get to and from Mars, along with a bunch of other stuff:

http://www.lpi.usra.edu/meetings/leag2007/presentations/20071001.drake.pdf

Note the Orion vehicle on the cover page illustration of the Mars stack.  How else would you propose we get off the surface of Earth, and then land back on Earth?

Unfortunately, on page two, under "purposes", they say:

Probably does not represent the way we will end up going to Mars

Which kinda makes one wonder what has been the purpose of all the study since 1991?  To study all the ways you can't go to Mars?

[baddux]

That statement doesn't make a lot of sense to me, since Orion seems to be an integral part of how they plan to get to and from Mars, along with a bunch of other stuff:

http://www.lpi.usra.edu/meetings/leag2007/presentations/20071001.drake.pdf

Note the Orion vehicle on the cover page illustration of the Mars stack.  How else would you propose we get off the surface of Earth, and then land back on Earth?

My point is that Orion is too heavy and complex to do that in actual mission, 25mt vs 4mt that can do the job. You don't need a toilet in a landing and ascent vehicle. And you definitely don't want to lift extra 20mt to leo for nothing when you want to minimize launches per mission. Not to mention all the budget problems it would bring.

I'm aware of that DRM5 plan, youtube link: /watch?v=uUBhn3_P3hU
but nobody thinks this is a realistic plan or ever was.

[sdsds]

When it rains lemons, make lemonade. Orion's capabilities -- like supporting its crew for 21 days and providing meaningful propulsion -- could be put to good use. It adds flexibility in where the crew makes rendezvous with the departure stack, and where it parts with the return stack.

[A_M_Swallow]

My point is that Orion is too heavy and complex to do that in actual mission, 25mt vs 4mt that can do the job. You don't need a toilet in a landing and ascent vehicle. And you definitely don't want to lift extra 20mt to leo for nothing when you want to minimize launches per mission. Not to mention all the budget problems it would bring.

I'm aware of that DRM5 plan, youtube link: /watch?v=uUBhn3_P3hU
but nobody thinks this is a realistic plan or ever was.

Instead of meeting the Mars Transfer Vehicle (MTV) at LEO the Orion could dock with the MTV at an Earth-Moon Lagrange (EML) point.  Getting from Earth to an EML will take several days so the toilet is likely to be used.  Orion's heat shield will work from EML.

The Orion could be used as a bridge (control room) for the MTV and fly it by wire.

[baddux]

Instead of meeting the Mars Transfer Vehicle (MTV) at LEO the Orion could dock with the MTV at an Earth-Moon Lagrange (EML) point. 

Yeah, you can find uses for Orion but I believe they are unrealistic with NASA's budgets. NASA is now developing exactly those vehicles, Orion and SLS, that they shouldn't if they are planning to go to Mars. They have other stuff that are specific to Mars mission (hab module, mars rover, ISRU plant + MAV etc.). Dragon could serve as takeoff + LAS + return + probably even Mars landing vehicle, and it already exists and is relatively low cost. HLV could be done as commercial HLV program similar to commercial crew. ULA, Spacex and ATK + EADS could be possible competitors. Then NASA would actually have the money to develop the rest of the needed stuff, now they are all 'wasted' to Orion and SLS.

This fits to Zubrin's house building parable that you cannot first buy parts and then deside what kind of house you build so that all parts will be used.

[muomega0]

When it rains lemons, make lemonade. Orion's capabilities -- like supporting its crew for 21 days and providing meaningful propulsion -- could be put to good use. It adds flexibility in where the crew makes rendezvous with the departure stack, and where it parts with the return stack.

Everyone wants to know how the lemonade is protecting the crew from galactic cosmic radiation at L2 and beyond, and if this weight is included in the design that returns to earth.

if the trip to L2 takes 3 days or less, why include gcr protection at all? include it in the habitat at L2.   Why take the weight back and forth?

If one is going to exclude the gcr protection, how can anyone on earth think they can claim a 21 day capacity for Orion BEO?!

Does anyone think of these things?

[JohnFornaro]

Note the Orion vehicle ...  How else would you propose we get off the surface of Earth, and then land back on Earth?

1) My point is that Orion is too heavy and complex to do that in actual mission, 25mt vs 4mt that can do the job. You don't need a toilet in a landing and ascent vehicle. And you definitely don't want to lift extra 20mt to leo for nothing when you want to minimize launches per mission. Not to mention all the budget problems it would bring.

2) I'm aware of that DRM5 plan, ... but nobody thinks this is a realistic plan or ever was.

True about the toilet, but the OML and mating ring are fine.  There is no doubt in my mind that there's plenty of commonsense commonality in the Orion capsule such that several different functions can be met within the same OML.  Furthermore, the extra mass and volume will provide margin for the long trip to Mars.

As to your assertion that 4 tons will do the job; this cannot be accepted without substantiation.  There would certainly be budget problems in developing such a clean sheet design which would provide the low mass you speculate on.

Backing up the toilet, uhhhh... backing up to the toilet issue, and capsule commonality.  Hopefully, over the coming years, NASA will realize the benefits of amortizing costs over a vehicle "line".  According to a lot of posters here, minor differences in vehicle functionality can and should only be optimized in one-off designs, necessitating not only name changes to the program, but reams of clean sheets, and the attendant budgetary bloat.  Which gets to Thing #2.

2) The DRM5 plan represents the efforts of our "best and brightest".  You and a host of others have opined about how unrealistic it is.  Even the program authors opine: "Probably does not represent the way we will end up going to Mars".

Unfortunately, the previous and current administrations, and the Congessi of those administrations, have not asked the important question:

What are you guys actually doing with the money you're receiving?

From the old eyes, small fonts department:

>

This fits to Zubrin's house building parable that you cannot first buy pants and then decide what kind of house you build so that all pants will be used.

What???

[baddux]

2) The DRM5 plan represents the efforts of our "best and brightest".  You and a host of others have opined about how unrealistic it is.  Even the program authors opine: "Probably does not represent the way we will end up going to Mars".

I don't doubt their expertise but I believe their assignment has been to design a Mars plan which uses (yet nonexisting) Ares V, Ares I and Orion even if you wouldn't do that if you could deside.

This fits to Zubrin's house building parable that you cannot first buy pants and then decide what kind of house you build so that all pants will be used.

What???

Yeah, it would be too bad if the mass requirements would be so tight that the astronauts would have to be without pants 

[sdsds]

Yeah, you can find uses for Orion but I believe they are unrealistic with NASA's budgets. NASA is now developing exactly those vehicles, Orion and SLS, that they shouldn't if they are planning to go to Mars.

This makes so much sense, except: it presupposes that NASA gets to choose where it spends its budget. That's just not the case. SLS is funded by Congress. Orion is funded by Congress. NASA is merely the agency chosen by Congress to carry out those projects.

When as a nation we ask, "How should we go about our human spaceflight effort" we tend to set up a blue-ribbon panel to give us answers. We did that with the Augustine Review committee. The Mars First community provided input just like everyone else. The conclusion was clear:

Mars is not a viable first destination beyond low-Earth or-
bit at this time. With existing technology and even a sub-
stantially increased budget, the attainment of even sym-
bolic missions would demand decades of investment and
carry considerable safety risk to humans. It is important
to develop better technology and gain more experience in
both free space and surface exploration prior to commit-
ting to a specific plan for human exploration of the surface
of Mars.

http://www.nasa.gov/pdf/396093main_HSF_Cmte_FinalReport.pdf

[baddux]

Especially suspicious in this DRM5 plan is that first the crew travels to Mars orbit in transit hab, then move with Orion to the Mars hab and then descent to the surface. This looks really risky thing to do in Mars orbit and makes you wonder if this just added so that they can use Orion to something other than takeoff and landing. Why don't they just travel with this surface hab like in Mars direct or DRM1.  If this will make the hab too heavy then use two habs with crew of 3 each.

[pathfinder_01]

Especially suspicious in this DRM5 plan is that first the crew travels to Mars orbit in transit hab, then move with Orion to the Mars hab and then descent to the surface. This looks really risky thing to do in Mars orbit and makes you wonder if this just added so that they can use Orion to something other than takeoff and landing. Why don't they just travel with this surface hab like in Mars direct or DRM1.  If this will make the hab too heavy then use two habs with crew of 3 each.

Nah, Zurbin assumes aerobraking(or aerocapture) for the crew followed by landing. Mars DRM 5.0 does not. Zurbin probably does not assume enough supplies and propellant on the Mars HAB to get home.

DRM 5.0 assumes that enough supplies to last the whole mission is carred in the crew transfer vechile so that if they don't land they don't run out of supplies and the transfer vechile does both inbound and outbound trips.
This means that this vechile is larger than what Zurbin would use.

Orion in this case is moving the crew between the two vechiles. The question is why not simply dock the CTV to the Habitat\lander? (i.e. Is the CTV in a higher orbit but the hab\lander in a lower one?)

[spectre9]

It's like Zubrin says.

NASA had to design a Mars mission that made use of everybody's pet technology.

DRA5 is that plan.

Any complexities observed are a result of this absolute mash of technology being forced into the mission.

[pathfinder_01]

Any complexities observed are a result of this absolute mash of technology being forced into the mission.

Perhaps. Something like Orion would be useful if you were using electric propulsion for the mission(or many deep space missions). You would build you mars transfer stack in LEO, then let it spiral out to a high earth orbit or L-point unmanned. This could take weeks or months.

You would use Orion to transfer crew from Earth or LEO to a high earth orbit or L-point.

You could use it again to depart the ship before it goes into orbit. Electric propulsion can take months to get into orbit but departing early would return the crew faster.  You could then reuse your mars transfer vehicle.

However in this case you would also need to carry a lander(or lander/hab) with you as you might depart the deep space craft before it goes into orbit around mars(again leaving it to go into mars orbit unmanned). 

However for mars drm 5.0 yeah there are some unrealistic and impractical points(like all mars plans). However the mass of Orion lets it have more delta V and more time in space(21 days total). It is just a question of does the mission use that mass and time or not.

[strangequark]

Especially suspicious in this DRM5 plan is that first the crew travels to Mars orbit in transit hab, then move with Orion to the Mars hab and then descent to the surface. This looks really risky thing to do in Mars orbit and makes you wonder if this just added so that they can use Orion to something other than takeoff and landing. Why don't they just travel with this surface hab like in Mars direct or DRM1.  If this will make the hab too heavy then use two habs with crew of 3 each.

Nah, Zurbin assumes aerobraking(or aerocapture) for the crew followed by landing. Mars DRM 5.0 does not. Zurbin probably does not assume enough supplies and propellant on the Mars HAB to get home.

DRM 5.0 assumes that enough supplies to last the whole mission is carred in the crew transfer vechile so that if they don't land they don't run out of supplies and the transfer vechile does both inbound and outbound trips.
This means that this vechile is larger than what Zurbin would use.

Orion in this case is moving the crew between the two vechiles. The question is why not simply dock the CTV to the Habitat\lander? (i.e. Is the CTV in a higher orbit but the hab\lander in a lower one?)

Zubrin's pretty skimpy with mass estimates in general. It's informative to compare some of his masses with actual flying spacecraft.

[baddux]

Zubrin's pretty skimpy with mass estimates in general. It's informative to compare some of his masses with actual flying spacecraft.

True. I think DRM1 is still the most viable Mars plan (Mars direct but return vehicle in mars orbit). It could actually be done with available budgets but the bad thing was that it required 240mt or something HLV with NASA's weight estimates. It had crew of 6 so I don't know if you would use two identical with crew of 3 what would be the HLV requirement then. In case of an emergency those could be docked in space.

[QuantumG]

True. I think DRM1 is still the most viable Mars plan (Mars direct but return vehicle in mars orbit). It could actually be done with available budgets but the bad thing was that it required 240mt or something HLV with NASA's weight estimates. It had crew of 6 so I don't know if you would use two identical with crew of 3 what would be the HLV requirement then. In case of an emergency those could be docked in space.

Or just do on-orbit assembly. It takes discipline to not turn your mission design into a battlestar galactica but there's gotta be a happy medium.

[baddux]

Especially suspicious in this DRM5 plan is that first the crew travels to Mars orbit in transit hab, then move with Orion to the Mars hab and then descent to the surface. This looks really risky thing to do in Mars orbit and makes you wonder if this just added so that they can use Orion to something other than takeoff and landing. Why don't they just travel with this surface hab like in Mars direct or DRM1.  If this will make the hab too heavy then use two habs with crew of 3 each.

Nah, Zurbin assumes aerobraking(or aerocapture) for the crew followed by landing. Mars DRM 5.0 does not. Zurbin probably does not assume enough supplies and propellant on the Mars HAB to get home.

DRM 5.0 assumes that enough supplies to last the whole mission is carred in the crew transfer vechile so that if they don't land they don't run out of supplies and the transfer vechile does both inbound and outbound trips.
This means that this vechile is larger than what Zurbin would use.

Right, I forgot this aerocapture thing. So they have extra consumables for 18 monts if they are not able to land to the surface? Quite a lot of extra weight for safety.

Also in Mars direct there was possibility to abort before Mars orbit, that's why Zubrin uses this 6 monts free return-trajectory afaik? Then the crew would travel 2 years so they would be in that hab the same time as in successful mission. That I don't know how they would then land to earth.

[Dalhousie]

True. I think DRM1 is still the most viable Mars plan (Mars direct but return vehicle in mars orbit). It could actually be done with available budgets but the bad thing was that it required 240mt or something HLV with NASA's weight estimates. It had crew of 6 so I don't know if you would use two identical with crew of 3 what would be the HLV requirement then. In case of an emergency those could be docked in space.

Personally I thought DRM 1.0 excessively large for what they did (much the same as 5.0).  DRM 3.0 and 4.0 were much more realistic, mass wise. IMHO!

[Dalhousie]

Nah, Zurbin assumes aerobraking(or aerocapture) for the crew followed by landing. Mars DRM 5.0 does not. Zurbin probably does not assume enough supplies and propellant on the Mars HAB to get home.

DRM 5.0 assumes that enough supplies to last the whole mission is carred in the crew transfer vechile so that if they don't land they don't run out of supplies and the transfer vechile does both inbound and outbound trips.
This means that this vechile is larger than what Zurbin would use.

Orion in this case is moving the crew between the two vechiles. The question is why not simply dock the CTV to the Habitat\lander? (i.e. Is the CTV in a higher orbit but the hab\lander in a lower one?)

If Mars Direct was unable to land they would be on an 800 day return trajectory.  The 18 months of supplies in the hab would be able to cover this eventuality.  The same applies to the initial Mars semi-direct approach, followed by the early DRMs, (1.0, 2.0 and, i think some versions of 3.0).

In DRMs 4.0 and 5.0 AFAIR the crew travels to and from Mars in a ferry, and the 18 month reserve has to be carried.  However, if unused, the reserved can be dumped before returning home, saving mass that way.

[baddux]

One good thing in this transhab concept is that it could also do a Mars orbit mission without modifications. Also it could be easily tested with incrementally harder missions, first just the hab in the ISS, then the whole vehicle in earth orbit, then in lunar orbit for whole Mars orbit mission time to test zero-g, radiation, life support etc.

So if NASA would cancel SLS and fund transhab instead and pick either NTP or chemical propulsion for EDS and develop that, then they could actually get closer to a Mars mission and meanwhile do some at least better than LEO mission (lunar orbit or EML1).

[Dalhousie]

One good thing in this transhab concept is that it could also do a Mars orbit mission without modifications. Also it could be easily tested with incrementally harder missions, first just the hab in the ISS, then the whole vehicle in earth orbit, then in lunar orbit for whole Mars orbit mission time to test zero-g, radiation, life support etc.

So if NASA would cancel SLS and fund transhab instead and pick either NTP or chemical propulsion for EDS and develop that, then they could actually get closer to a Mars mission and meanwhile do some at least better than LEO mission (lunar orbit or EML1).

Any Mars mission Transhab would need to be specifically fitted out .  It's no different to any other module in this respect.  Zarya or Zvezda or Columbus could just as easily be adapted for a Mars mission.  There is nothing about Transhab that makes it especially suitable.

[pathfinder_01]

Any Mars mission Transhab would need to be specifically fitted out .  It's no different to any other module in this respect.  Zarya or Zvezda or Columbus could just as easily be adapted for a Mars mission.  There is nothing about Transhab that makes it especially suitable.

Transhab is inflatable which in theory saves mass over Zarya/Zvezda and Colubus. However there is less data about inflatable structures in space than metal canisters. In terms of radation protection both are iffy.

[Dalhousie]

Transhab is inflatable which in theory saves mass over Zarya/Zvezda and Colubus. However there is less data about inflatable structures in space than metal canisters. In terms of radation protection both are iffy.

I am not convinced that Transhab saves much mass.  Most of the mass of a mission module comes from systems, not the pressure vessel.

Additional radiation protection, mostly likely a storm shelter, is one of the many modifications that would be needed for any module that ise based on an existing design.

[baddux]

One good thing in this transhab concept is that it could also do a Mars orbit mission without modifications. Also it could be easily tested with incrementally harder missions, first just the hab in the ISS, then the whole vehicle in earth orbit, then in lunar orbit for whole Mars orbit mission time to test zero-g, radiation, life support etc.

So if NASA would cancel SLS and fund transhab instead and pick either NTP or chemical propulsion for EDS and develop that, then they could actually get closer to a Mars mission and meanwhile do some at least better than LEO mission (lunar orbit or EML1).

Any Mars mission Transhab would need to be specifically fitted out .  It's no different to any other module in this respect.  Zarya or Zvezda or Columbus could just as easily be adapted for a Mars mission.  There is nothing about Transhab that makes it especially suitable.

Yes, I meant more the concept of the mission than the actual vehicle type. If the vehicle is designed to go to Mars orbit, stay operational for 18 months and come back to earth (probably staying in earth orbit or EML1 for the next mission), then it would be easy to do orbit only mission first and avoid the need for landing vehicle and have some smaller MAV for retrieving samples robotically to the orbiting hab, it could also test ISRU.

[Dalhousie]

Yes, I meant more the concept of the mission than the actual vehicle type. If the vehicle is designed to go to Mars orbit, stay operational for 18 months and come back to earth (probably staying in earth orbit or EML1 for the next mission), then it would be easy to do orbit only mission first and avoid the need for landing vehicle and have some smaller MAV for retrieving samples robotically to the orbiting hab, it could also test ISRU.

Zubrin looked into this with the Athena concept, http://www.astronautix.com/craft/athena.htm as have quite a few others, of course.

These studies look at mision requires - crew size, volume, mass, consumables, propulsion etc, which fine.  The rouble is that as far as I know, nobody has looked in detail and the surface mission, what would actually be required, what would actually be achieved, etc. To me at any rate the surface mission is what it's all about.  So I would be delighted if I was pointed to studies out there detailing this.

[enginerd]

Does anyone know where I can find Zubrin's Mars Direct paper that has all of the figures in it? Seems to me that every file online is missing most, if not all of the figures. I'm trying to play catch up on this subject....

[Kaputnik]

I think you have to buy his book. You can probably get a cheap used copy on Amazon or eBay as it has been around for quite a few years now.
(I got mine from the local library so I don't have my own copy to look up)

[truth is life]

I think you have to buy his book. You can probably get a cheap used copy on Amazon or eBay as it has been around for quite a few years now.
(I got mine from the local library so I don't have my own copy to look up)

However, he still tends to be vague at best about where his numbers come from, tending to appeal to "reasonable engineering estimates" (ie., his own authority) or the like instead of saying something like, "a heat shield to enter Mars' atmosphere directly from a trans-Mars trajectory will weigh xxx pounds per square inch based on current technology, and the maximum ballistic coefficient we can handle and still land safely is yyy pounds per square inch. Adding in our 20 ton habitat/return vehicle, and this means we need a heat shield zzz feet across which will weight aaa pounds for aerobraking". There might be good reasons for this in the main narrative, but he does this in the appendices and technical notes as well...

[KelvinZero]

However, he still tends to be vague at best about where his numbers come from,

Perhaps it would be worth looking at "mars semi direct". Less pure Zubrin but probably more numbers?
(Just a guess, Ive never looked at it and am not an engineer. I understand it was produced by NASA though)

[Dalhousie]

However, he still tends to be vague at best about where his numbers come from,

Perhaps it would be worth looking at "mars semi direct". Less pure Zubrin but probably more numbers?
(Just a guess, Ive never looked at it and am not an engineer. I understand it was produced by NASA though)

Zubrin and Weaver came up with semi-direct as well, as a result of feedback over MD.  It has been the archiecture of choice for most studies since.

[Dalhousie]

However, he still tends to be vague at best about where his numbers come from, tending to appeal to "reasonable engineering estimates" (ie., his own authority) or the like instead of saying something like, "a heat shield to enter Mars' atmosphere directly from a trans-Mars trajectory will weigh xxx pounds per square inch based on current technology, and the maximum ballistic coefficient we can handle and still land safely is yyy pounds per square inch. Adding in our 20 ton habitat/return vehicle, and this means we need a heat shield zzz feet across which will weight aaa pounds for aerobraking". There might be good reasons for this in the main narrative, but he does this in the appendices and technical notes as well...

It did not need that level of detail for the purpose for which it was intended.

[QuantumG]

It did not need that level of detail for the purpose for which it was intended.

Indeed. An analogy would be someone drawing a conceptual design of a building and receiving criticism for failing to explain where the air conditioning would go.

[go4mars]

At 24:40 or so, he talks about an international effort in the spirit of the Olympics.  A friendly, competitive basis. 

Anyone have some ideas about what form that might take?  Or a link to somewhere that Bob Zubrin goes into more detail about it?

[go4mars]

This politician, Adam Schiff gives an exciting address.  Very pro Mars exploration.

[truth is life]

It did not need that level of detail for the purpose for which it was intended.

This is true, which I why I noted that there were probably reasons for this in the main narrative. However, it shows up in the ostensibly technical notes as well, which bothered me. Of course, this may just be me grousing because I just went back and looked at it again for a role in which the extra technical detail would have been helpful, only I didn't find it You can't go look up Bob Zubrin's intuition on NTRS, after all...

[Robotbeat]

6 MW gets the round trip down to 13 months which is within the limits of what is known as far as mission duration goes.

Do you have any idea what the largest space nuclear power source ever flown is?

My masses were for solar not nuclear.

Fine.. do you have any idea what the biggest solar power source ever flown is? (and the power to mass of that system?)

If you believe in solar powered plasma rockets, you might as well believe in solar power satellites too. Zubrin calls it "an alternate universe".

Solar powered plasma rocket? No one says you HAVE to use VASIMR. There are lots of other electric thrusters... Roughly half of active comm-sats at GSO use some form of electric thrust.

The largest solar power source ever flown is the ISS, about 110kW.... Far higher than the highest power nuclear power source ever flown. ISS is a very poor example of specific power, though, since it is very heavy and uses quite old technology. Current state of the art for specific power is the UltraFlex arrays of Phoenix, Orion, etc, which can do up to 150W/kg or more (more for larger arrays).

At that specific power, the ISS's arrays would weigh less than a ton. A useful tug would need at LEAST that much power... If you were to design a system clean-sheet, I'd probably aim for 350-500kW, which is quite capable and could be used for pushing stuff all over the inner solar system (and even provide plenty of useful power at Jupiter and Saturn, if you were careful).

A 500kW array would need to be no larger than the ISS's array (in total length and width), but would be much more efficient and would have a coverage ratio much higher. A 500kW array would have a mass of around 2-4 metric tons, if using similar technology (or equivalent performance) as UltraFlex (the proposal for using UltraFlex in the hundreds of kilowatts up to a megawatt is called MegaFlex http://sbir.gsfc.nasa.gov/SBIR/abstracts/11/sbir/phase1/SBIR-11-1-X8.04-9203.html?solicitationId=SBIR_11_P1 ... expects to hit 200W/kg specific power).

There are other SBIR proposals for >500W/kg and even some for up to >1000W/kg... That puts even a 6MW solar array at just 6 or 12 metric tons, could be launched in one go by several launch vehicles. For an integrated propulsion unit in the middle Megawatts, you might need to use an EELV Heavy class launch vehicle even if you got 500-1000W/kg specific power. But there is no fundamental reason why those specific powers couldn't be reached. It will take more research, though.

But we can definitely develop a useful tug using current technology and launchable on existing launch vehicles.

[GalacticIntruder]

Space gymnastics, you must stick the landing to get the gold.

After watching all the MSL coverage, it hit me, that landing anything on the surface of Mars is freaking difficult. Only the US has done it. All these Mars missions just sort assume you send large amounts of mass accurately to Mars' surface; and sometimes many times, ie Mars One. I am not aware anyone is even working on that problem for human missions, much less repeatable, affordable, and timely. True, SpX has Red Dragon, but I am not convinced they thought it through in extreme detail. Orion cannot land either.  Unless someone can make something work, I cannot see any way to send humans, a crew of two or six, for 18 months.

Anyone know anything tangible?

[Kaputnik]

After watching all the MSL coverage, it hit me, that landing anything on the surface of Mars is freaking difficult. Only the US has done it.

Erm, not quite. The Soviets landed the first spacecraft on Mars, although it stopped functioning shortly after. But yes, it is bloomin' difficult to land anything on Mars.

I am not aware anyone is even working on that problem for human missions

It is acknowledged as one of, if not the single biggest, difficulties that needs to be overcome before a human mission can take place. NASA are currently working on a prototype inflatable heatshield technology which could certainly help. It would be good to see more emphasis in this area, but any development is likely to be very expensive. That's why NASA is still using Viking technology- they spent a lot of money back in the 70s and don't want to have to do it again.

True, SpX has Red Dragon, but I am not convinced they thought it through in extreme detail. Orion cannot land either.

Red Dragon may work, but IMHO would be a bit of a stunt. A more capable lander could be designed using SpaceX technologies that would be more suited to the task.
Orion is not, and never has been, and never will be, designed to land on Mars. Mars really needs a dedicated lander, just like the moon or Earth also need dedicated vehicles.

[JohnFornaro]

It did not need that level of detail for the purpose for which it was intended.

Indeed. An analogy would be someone drawing a conceptual design of a building and receiving criticism for failing to explain where the air conditioning would go.

Both of you are right, but if you want to get the building built per your design vision, you will have to create detailed architectural plans, structural and mechanical engineering drawings, and complete specifications.

It seems like Mr. Zubrin needs to get to the drawing board and tighten up the design.

So where exactly, is that air conditioning system?

[QuantumG]

No.. the detailed design is something a whole team does.. not an advocate asking why the whole kitchen sink has to be included in every exploration plan.

There was basically one technology which defined the 90 day study - the Battlestar Galactica plan that Zubrin and co were rallying against: aerobraking into LEO. The only reason why this new technology was considered "critical" was so Space Station Freedom, and the Shuttle, could have a role both at the start and at the end of the exploration mission.

Today we have different "critical technologies", but it's essentially the same argument: we're not ready to explore the solar system with humans yet.

I don't get behind much that Zubrin says, but his insistence that we could design exploration architectures with existing technology if we really wanted to go start doing exploration right now, is right on the money. The fact that we keep looking for the next "critical" technology before exploration can start is just proof that we're not all that interested in going anywhere.

[Solman]

6 MW gets the round trip down to 13 months which is within the limits of what is known as far as mission duration goes.

Do you have any idea what the largest space nuclear power source ever flown is?

My masses were for solar not nuclear.

Fine.. do you have any idea what the biggest solar power source ever flown is? (and the power to mass of that system?)

If you believe in solar powered plasma rockets, you might as well believe in solar power satellites too. Zubrin calls it "an alternate universe".

Solar powered plasma rocket? No one says you HAVE to use VASIMR. There are lots of other electric thrusters... Roughly half of active comm-sats at GSO use some form of electric thrust.

The largest solar power source ever flown is the ISS, about 110kW.... Far higher than the highest power nuclear power source ever flown. ISS is a very poor example of specific power, though, since it is very heavy and uses quite old technology. Current state of the art for specific power is the UltraFlex arrays of Phoenix, Orion, etc, which can do up to 150W/kg or more (more for larger arrays).

At that specific power, the ISS's arrays would weigh less than a ton. A useful tug would need at LEAST that much power... If you were to design a system clean-sheet, I'd probably aim for 350-500kW, which is quite capable and could be used for pushing stuff all over the inner solar system (and even provide plenty of useful power at Jupiter and Saturn, if you were careful).

A 500kW array would need to be no larger than the ISS's array (in total length and width), but would be much more efficient and would have a coverage ratio much higher. A 500kW array would have a mass of around 2-4 metric tons, if using similar technology (or equivalent performance) as UltraFlex (the proposal for using UltraFlex in the hundreds of kilowatts up to a megawatt is called MegaFlex http://sbir.gsfc.nasa.gov/SBIR/abstracts/11/sbir/phase1/SBIR-11-1-X8.04-9203.html?solicitationId=SBIR_11_P1 ... expects to hit 200W/kg specific power).

There are other SBIR proposals for >500W/kg and even some for up to >1000W/kg... That puts even a 6MW solar array at just 6 or 12 metric tons, could be launched in one go by several launch vehicles. For an integrated propulsion unit in the middle Megawatts, you might need to use an EELV Heavy class launch vehicle even if you got 500-1000W/kg specific power. But there is no fundamental reason why those specific powers couldn't be reached. It will take more research, though.

But we can definitely develop a useful tug using current technology and launchable on existing launch vehicles.

 I've been looking through Spectrolab data sheets in an unsuccessful attempt to find the specific power of their cells for 500X or thereabouts. Efficiencies of nearly 40% at over 900X solar concentration factor compared to solar intensity at 1 AU are cited and on page 67 of this sheet: http://www.spectrolab.com/DataSheets/PV/pv_tech/msce.pdf
they claim a factor of 1000 advantage in area compared to 17% cells with 500X cells at 35%. It seems to me that if you are just talking about the mass of the solar cells and not the total mass of the system that their specific power would be in the 500X ballpark even vs. say 30% eff. type that does not use concentrated sunlight, provided the mass per unit area of the concentrator type is about the same. It seems unlikely that they are more than a few times as massive per area. If they are even 5X as heavy per area they would produce 100X times the power per unit mass. If the 30% type are 150W/kg then the concentrator cells would seem to be 15000W/kg.
 Now admittedly I could not find this explicitly stated in the data sheets, but it sure looks logical to me. If this is true then using the 17KW/kg L'Garde concentrator would result in an overall system specific power in the 7.5KW/kg range if my often flawed logic is correct.

That would be absolutely game changing!

[Robotbeat]

The mass of the concentrator is not small. I used to be a big fan of concentrator systems, but you can get comparable specific powers (or even better) by using efficient thin-film techniques. Concentrators have had problems in the past.


Though they are a pretty good solution for outer solar system missions, since low-light efficiency is disproportionately less, so increasing the intensity helps a lot.

[QuantumG]

That would be absolutely game changing!

Before something can be "game changing" there has to be a "game".

Currently the only game is who can come up with the next "critical" technology or capability which must be developed before exploration can begin.

[MATTBLAK]

No.. the detailed design is something a whole team does.. not an advocate asking why the whole kitchen sink has to be included in every exploration plan.

There was basically one technology which defined the 90 day study - the Battlestar Galactica plan that Zubrin and co were rallying against: aerobraking into LEO. The only reason why this new technology was considered "critical" was so Space Station Freedom, and the Shuttle, could have a role both at the start and at the end of the exploration mission.

Today we have different "critical technologies", but it's essentially the same argument: we're not ready to explore the solar system with humans yet.

I don't get behind much that Zubrin says, but his insistence that we could design exploration architectures with existing technology if we really wanted to go start doing exploration right now, is right on the money. The fact that we keep looking for the next "critical" technology before exploration can start is just proof that we're not all that interested in going anywhere.

Exactly so, Quantum!! If new technology keeps being put on the critical path to achieve exploration missions, then permanent deferral will happen. Lunar missions scarcely need any new tech to achieve missions, though without it, there might only be 'Apollo on Steroids' for the interim. Which would be fine by me! But to do better than 'Steroids', develop a re-usable lander and some form of Propellant Depot. A hypergolic depot at L-1, topped up by a Commercial propellant 'supply line' would enable a regular string of long 'Sortie' missions that could lead to a man-tended, semi-permanent Lunar Outpost. LOX/LH2 from the Lunar Poles could come later: putting cryogenic ISRU in the critical path now will cost a lot and slow things down.

But in the long term, it will bring benefits. As for Heavy Lift: it costs a lot up front in the short term but makes certain mission architectures easier. But in the long term, it may be unsustainable. For Mars missions, nuclear-powered ISRU could be viewed as highly desirable, but perhaps not essential - do it with advanced solar power, leveraged by Stirling engine technology and have smaller crews needing less power and equipment. Same with Heavy Lift for Mars: Even with SLS Block II you will need a lot of launches. With EELV lift; yes, a lot more launches of even uprated EELVs.

But by using the entire International rocket 'fleet' launched in multiple Salvos and leveraged with Propellant Depots and Solar Electric Propulsion as well - it could be done. Any way we slice it, it's going to be difficult and expensive. But for the coming future we have to find ways to do Exploration smarter and not with 'brute force' methods that Congresses, Presidents and Parliaments simply wont pay for.

[QuantumG]

Exactly so, Qunatum!! If new technology keeps being put on the critical path to achieve exploration missions, then permanent deferral will happen. Lunar missions scarcely need any new tech to achieve missions, though without it, there might only be 'Apollo on Steroids' for the interim. Which would be fine by me! But to do better than 'Steroids', develop a re-usable lander and some form of Propellant Depot.

With you so far...

A hypergolic depot at L-1, topped up by a Commercial propellant 'supply line' would enable a regular string of long 'Sortie' missions that could lead to a man-tended, semi-permanent Lunar Outpost. LOX/LH2 from the Lunar Poles could come later: putting cryogenic ISRU in the critical path now will cost a lot and slow things down.

Why a hypergolic depot? Every vehicle that has flown in space in the last 30 years has used a particular propellant combination and it isn't hypergolic. Why start with technology almost no-one is using - that's backwards.

[MATTBLAK]

Actually, I'd prefer LOX/Kerosene! Or LOX/Methane... But the Isp of hypergolic isn't too bad compared to hydrocarbon propellants. As I said, putting LOX/LH2 on the critical path will slow things and in the short term cost a lot to perfect. As for hypergolic; excellent for long-term storage, a mature technology used in a lot of spacecraft (ISS, Shuttle, Dragon, Apollo) probes and satellites. Also, it has proven pumping and transfer methods and is simple. mmeijeri is better at outlining these things than I.

[Robotbeat]

Exactly so, Qunatum!! If new technology keeps being put on the critical path to achieve exploration missions, then permanent deferral will happen. Lunar missions scarcely need any new tech to achieve missions, though without it, there might only be 'Apollo on Steroids' for the interim. Which would be fine by me! But to do better than 'Steroids', develop a re-usable lander and some form of Propellant Depot.

With you so far...

A hypergolic depot at L-1, topped up by a Commercial propellant 'supply line' would enable a regular string of long 'Sortie' missions that could lead to a man-tended, semi-permanent Lunar Outpost. LOX/LH2 from the Lunar Poles could come later: putting cryogenic ISRU in the critical path now will cost a lot and slow things down.

Why a hypergolic depot? Every vehicle that has flown in space in the last 30 years has used a particular propellant combination and it isn't hypergolic. Why start with technology almost no-one is using - that's backwards.

Huh? Almost every spacecraft since the dawn of the space age has been using hypergolic propellants especially for in-space maneuvering, almost always hydrazine-based. The ISS is essentially a hypergolic propellant depot, with propellant delivered via ATV and Progress.

[MATTBLAK]

Yes.

[KelvinZero]

Surely we don't need or want to put a reusable lander on the critical path either though. Reusable can delay getting started and then slow evolution. How about start with an minimalist expendable approach with a general strategy of how to evolve to reusable later, if volume justifies it? After all you could reuse a launcher by dismantling it on the surface or keeping it as a backup. No nightmare of a crew landed with no vehicle that could get them home.

Im a "Lander firster" if there is such a thing. The HLV, Ascent stage (or method if reusable) and even people can wait.

So long as the majority of HSF money is going to launcher and ISS-style operations, every time you ask the professionals what is the next logical step they are going to keep coming back with nicely informed arguments that the best we can do next is something that involves a lot of launchers and ISS-style operations.

Im convinced that if we could just start landing infrastructure, then all this exploration technology and ISRU research would cease to be just toys waved at the public occasionally to justify more launcher investment. It would become big business and become a powerful lobby calling for more actual equipment on the surface actually doing something. The new mix of professionals would be no more or less intelligent than the old, but I predict they would start advocating totally different things as possible and reasonable.

There is very little that needs to be put on the critical path between here and landing HSF-relevant infrastructure on the moon. We could start landing human-scaled rovers with only a few hundred kg for example. The only reason we cannot do that AND any other HSF mission you can name is that we keep choosing HSF missions that are clearly too expensive even by themselves, and pretend they might work if we kill everything else to fund them.

[MATTBLAK]

I only mentioned a re-usable Lander mainly to appease the 'we don't wanna no Apollo-on-steroids' folk who always pipe up and mention how wasteful it is to throw-away hardware each time. I also said that I'd be happy with "Apollo 'Roids" style missions. You're right that putting a re-usable Lander in the critical path might slow things. But if we made the 'Block 1' Lander disposable with later 'Block 2' and beyond upgradeable to re-usability, that could be worth pursuing - whether they are fueled by hypergolics or cryogenics.

Though some could say that doing it re-usable from the word go would in the long term save money and effort and they maybe right. Though for Mars missions I couldn't see a re-usable Ascent/Descent Module happening as a critical-path item.

[spectre9]

There seems to be the promise that once the SLS/Orion development is over that these systems will then be cheap and affordable for exploration missions.

As long as that doesn't look likely there's going to be heavy scepticism about NASA's plans to go BEO.

They know exactly what they want, DSH and/or Lander.

If they don't get it then SLS/Orion are useless.

Gateway is just to make SLS/Orion look useful without them. Not many are falling for it. 

I can see many programs being wound up before 2020 such as JWST and the funding being given for a heavy lift mission architecture.

I like big volume launchers but it's easy to rag on them without a payload.

[A_M_Swallow]

Surely we don't need or want to put a reusable lander on the critical path either though. Reusable can delay getting started and then slow evolution. How about start with an minimalist expendable approach with a general strategy of how to evolve to reusable later, if volume justifies it? After all you could reuse a launcher by dismantling it on the surface or keeping it as a backup. No nightmare of a crew landed with no vehicle that could get them home.
{snip}

Cargo can be delivered to the Moon using expendable landers but people need to return to orbit.

On the Moon a reusable lander is basically a refuelable lander with big fuel tanks.  The pre-prototype landers are already refuelable on the Earth.  So the big difference is size.

If we can get the cabin including docking port, air lock, avionics, structure, astronauts and life support down to say 1 tonne then the specifications of the propulsion system become:

payload: 1 tonne
delta-V (return EML-1/2 to lunar surface): 2 * 2.52 = 5.04 km/s
lunar gravity: 1.622 m/s²
air restartable
in-space refuelable
able to take 14 days in the sun
throttle: 10:1
propellants: negotiable including possible inert flush

p.s. the SEV (Space Exploration Vehicle concept) including people weights 4000 kg.
http://www.nasa.gov/pdf/464826main_SEV_FactSheet_508.pdf

[Robotbeat]

Assume we decide to have a cislunar gateway and we stage our missions from there. What is the difference between an expendable and partially reusable lander if they are both staged from the same place, either a Lagrange gateway or one in polar orbit? List all differences.

[HappyMartian]

Exactly so, Qunatum!! If new technology keeps being put on the critical path to achieve exploration missions, then permanent deferral will happen. Lunar missions scarcely need any new tech to achieve missions, though without it, there might only be 'Apollo on Steroids' for the interim. Which would be fine by me! But to do better than 'Steroids', develop a re-usable lander and some form of Propellant Depot.

With you so far...

A hypergolic depot at L-1, topped up by a Commercial propellant 'supply line' would enable a regular string of long 'Sortie' missions that could lead to a man-tended, semi-permanent Lunar Outpost. LOX/LH2 from the Lunar Poles could come later: putting cryogenic ISRU in the critical path now will cost a lot and slow things down.

Why a hypergolic depot? Every vehicle that has flown in space in the last 30 years has used a particular propellant combination and it isn't hypergolic. Why start with technology almost no-one is using - that's backwards.

Huh? Almost every spacecraft since the dawn of the space age has been using hypergolic propellants especially for in-space maneuvering, almost always hydrazine-based. The ISS is essentially a hypergolic propellant depot, with propellant delivered via ATV and Progress.

If we are talking about reducing the mission costs of a reusable lander, then every pound that you don't have to land on the Moon reduces the Launcher weight by 500 pounds.

Using the most efficient propellant combination possible is the only way to reduce launcher requirements. Hydrolox is the most efficient propellant.

Hydrolox provides the most thrust for your money on the surface of the Moon. We have decades of experience in using hydrolox in space. Highly efficient hydrolox was the propellant combination for the Saturn V 2nd and 3rd stages. And the propellant burned by the Space Shuttle rocket engines. Hydrolox is also going to be used by the SLS engines.

Hydrolox was also the propellant of choice for the descent stage of NASA's proposed Altair Lunar Lander's descent stage that would have also made the the lunar orbit insertion (LOI) burn for the Orion Altair stack.

"The Altair descent stage was to be powered by four RL-10 rocket engines that are used for the Centaur upper stage used on the Atlas V rocket. Unlike the current RL-10 engines in use, the newer RL-10s would be able to throttle down to as low as 10% rated thrust (the current specifications allow for 20%), thus allowing the use of Altair for both the lunar orbit insertion (LOI) and landing stages of the lunar mission."

From: Constellation program
At: http://en.wikipedia.org/wiki/Project_Constellation#Altair



"Mission requirements obliged the vehicle to be able to descend from an equatorial or high-inclination lunar orbit to a polar landing site, along with bringing it and the Orion spacecraft into lunar orbit, as the Orion spacecraft's onboard Aerojet AJ-10 rocket engine and the amount of fuel it carries are insufficient to brake the Orion/Altair stack into lunar orbit (also crucial if it was flown unmanned without an Orion crew). The new lander would have been powered by a modified RL-10 engine (currently in use on the upper stage of the Delta IV rocket and Centaur upper stage of the Atlas V rocket), burning liquid hydrogen (LH2) and liquid oxygen (LOX) for the descent phase. A single AJ-10 rocket engine, like that on the Orion, was intended to power the ascent stage."

And, "The development of Altair would have been managed by the Constellation Lunar Lander Project Office at Johnson Space Center (JSC). JSC worked directly with Apollo astronauts, various industry suppliers and universities to develop the architecture for Altair. In conjunction with early development a mockup or testbed was to have been developed at JSC to study/develop specialized subsystems and other design considerations. Northrop Grumman, which built the Apollo Lunar Module, was contracted to help the project office develop the system concept."

From: Altair (spacecraft)
At: http://en.wikipedia.org/wiki/Altair_%28spacecraft%29



Constellation's  Altair hydrolox descent stage design also offers us what could be a reusable hydrolox cargo LM lander with large downmass, and potentially even a LEO to Lunar orbit reusable hydrolox tanker.

"The 53,600 kg cargo lander concept deletes the ascent stage and replaces it with 14,631 kg of cargo on the upper deck. Cargo capacity can be increased depending on the landing site. The requirement is 15-17 metric tons of dedicated cargo anywhere on the surface."

From: Altair project buying into Orion lessons for development process
February 29, 2008 by David Harris, Chris Bergin
At: http://www.nasaspaceflight.com/2008/02/altair-project-buying-into-orion-lessons-for-development-process/



The Altair descent stage based cago Lander LM offers "14,631 kg of cargo", which could include some of the equipment needed for a small scale Lunar ice ISRU hydrolox production facility. With four such Altair cargo LM landings, you could have the hydrolox production equipment you need to start up your ISRU facility.

Note also:


"The Lunar sortie mission, the focus of this report, provides global Lunar landing site access for short surface stay times along with early return capability. Current mission planning includes a maximum 7-day surface stay. The Lunar sortie technique provides a minimized worst-case CEV on-orbit plane change capability that does not exceed a specified value for any sortie landing site location."

"The Lunar polar technique focuses on landing sites in the Lunar polar regions greater than 85° latitude or less than –85° latitude. A polar sortie mission covers a maximum 7-day surface stay, while an outpost or long-duration mission covers a surface stay on the order of 180 days plus a 30-day contingency stay, for a total possible surface stay of 210 days."

From: Lunar Orbit Insertion Targeting and Associated Outbound Mission Design for Lunar Sortie Missions
At: http://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/20070025021_2007024501.pdf


If you want to save money for the American taxpayers, and taxpayers around the world, Lunar produced hydrolox is the only logical propellant for the reusable LM.

If we space cadets want to remain babies that can only take baby steps and are highly dependent on forever sucking the honey sweet money milk produced by taxpayers, we will simply stay far away from Lunar hydrolox and instead regularly propose other propellants and space exploration architectures that don't involve getting Lunar ice ISRU hydrolox production going as soon as possible and propose more dead end plans to keep us as space exploration infants going nowhere.

If space cadets really want to grow up soon and have some adult independence, we have no choice but devise a practical and doable space architecture that is based on Lunar ISRU hydrolox production and can soon enable reusable LM Landers and beyond cislunar spaceships and dramatically reduce spaceflight costs.


Cheers!


Edited.

[MATTBLAK]

Thanks for the links and posting(s), Happy Martian. But over the last 5 or 6 years, most of us here have already read thousands of pages of 'ESAS', Constellation and other Exploration program documents. At least I have - just how much of the arcane detail I have retained: I cannot say at this point...

And many of the 'Altair' technical details, equipment and proposed epic capabilities were sadly rendered academic by history; neutered by poor support and funding before it even had a chance for development.

But again; most of us know all that already...

[HappyMartian]

Assume we decide to have a cislunar gateway and we stage our missions from there. What is the difference between an expendable and partially reusable lander if they are both staged from the same place, either a Lagrange gateway or one in polar orbit? List all differences.

Initially stage from Lunar polar orbit, but as quickly as possible, switch to a polar region where the ISRU equipment and  storage propellant tanks are. The reusable LMs can be covered or "housed" on the Moon in movable garages in natural or robot made ultra cold traps.

The Altair design gives a potential large reusable LM cargo lander and LEO to to Lunar polar orbit tanker spacecraft. Note that without the hydrolox tanks and engines this basic structure and guidance system and software might also serve for a cargo module for a solar powered tug.

What is needed first is a minimalistic lightweight reusable hydrolox lander design that is small. How small is the question. For a WAG, assume a no cabin and open jeep platform, a crew of two or three wearing space suits, and precise navigation systems that allow quick landing and return missions that last less than an hour.

The lightweight lander also has to be usable by robotic rovers and automatically flyable without any crew on orbital and non orbital research missions. Perhaps simply a smaller version of the Altair cargo carrier would allow the same navigation computers and software to be used on both vehicles.

Horizontal LM landers are nice. Could the Altair cargo lander design be converted to a horizontal LM Lander? 


 

Edited.

[HappyMartian]

But again; most of us know all that already...

I know. But sometimes new folks show up. 

 

[QuantumG]

Huh? Almost every spacecraft since the dawn of the space age has been using hypergolic propellants especially for in-space maneuvering, almost always hydrazine-based. The ISS is essentially a hypergolic propellant depot, with propellant delivered via ATV and Progress.

Ha! That's what I get for posting at 4am while listening to The Space Show. I was sure he said cryogenic.... please disregard my sleep typing.

[HappyMartian]

....
able to take 14 days in the sun
....
propellants: negotiable including possible inert flush

The propellant is not negotiable because it isn't a political issue and making it a political issue could easily destroy this small scale robot based doable Lunar ice ISRU propellant project. This initial ISRU based Lunar exploration architecture would focus on making and using hydrolox unless an international scientific and engineering consensus emerges that meets all three of the following criteria:
1. There is a more efficient propellant combination that could be more easily made in the polar regions of the Moon from readily available ice and volatile gas resources.
2. That we currently have enough experience with that substitute propellant combination in rockets to use the substitute propellants as the basis of of the Lunar polar exploration and exploitation architecture.     
3. The use of such substitute propellants would allow less mass being lifted from the Earth than would the use of hydrolox in this initial robot ISRU project.

Other propellants might play a large role once we have established a permanent beachhead in a Lunar polar area, but right now we need to focus on what is the most doable and high efficiency propellant that we could readily make on the Lunar surface. Historically, that propellant is hydrolox.

'Making' hydrogen and oxygen out of water is so simple that even a child can do it. Compressing and turning the hydrogen and oxygen gasses into liquids is a more difficult task, but the natural and robot made ultra cold traps, and readily available regolith for insulation, would help to minimize the size or weight of the needed compressors, radiators, various storage tanks, and pipes.

As for your "able to take 14 days in the sun" there is no need for such a requirement for two clear reasons.
1. The polar regions have large areas with permanent or semi-permanent shadowing from the Sun. 
2. In Nineveh, one of the world's oldest cities, many thousands of years ago someone used a parasol to block sunlight. Eventually, an enormous Mylar parasol or umbrella could be erected to protect a LM Lander from solar heat. Berms and craters might also be used at some non polar locations to provide both shade and some protection from radiation.

Please be careful not to burden the initial minimalist lightweight Lunar hydrolox Landers that will operate only in the Lunar polar regions and in space with any unneeded requirements. We need a bare bones reusable LM Lander, not a universal go everywhere at anytime of the day type of LM lander. That will come later as modifications to the initial LM Landers or as new LM Lander designs that possibly use other propellants.


Cheers!

Edited.

[MATTBLAK]

Huh? Almost every spacecraft since the dawn of the space age has been using hypergolic propellants especially for in-space maneuvering, almost always hydrazine-based. The ISS is essentially a hypergolic propellant depot, with propellant delivered via ATV and Progress.

Ha! That's what I get for posting at 4am while listening to The Space Show. I was sure he said cryogenic.... please disregard my sleep typing.

Heh heh! I was up to 3am here with a stomach bug - at least you were having more fun than I was!

[KelvinZero]

Surely we don't need or want to put a reusable lander on the critical path either though. Reusable can delay getting started and then slow evolution. How about start with an minimalist expendable approach with a general strategy of how to evolve to reusable later, if volume justifies it? After all you could reuse a launcher by dismantling it on the surface or keeping it as a backup. No nightmare of a crew landed with no vehicle that could get them home.
{snip}

Cargo can be delivered to the Moon using expendable landers but people need to return to orbit.

On the Moon a reusable lander is basically a refuelable lander with big fuel tanks.  The pre-prototype landers are already refuelable on the Earth.  So the big difference is size.

If we can get the cabin including docking port, air lock, avionics, structure, astronauts and life support down to say 1 tonne then the specifications of the propulsion system become:

payload: 1 tonne
delta-V (return EML-1/2 to lunar surface): 2 * 2.52 = 5.04 km/s
lunar gravity: 1.622 m/s²
air restartable
in-space refuelable
able to take 14 days in the sun
throttle: 10:1
propellants: negotiable including possible inert flush

Air startable? 

Im glossing over the restartable detail, but a one way lander seems a good basis for a two way descent/ascent vehicle if fuel can be produced on the surface. I just mean on the moon, without air, descent and ascent are pretty similar I guess? If it can land from orbit, it can be refueled and ascend to orbit? Admittedly you might not have abort to orbit.
It is so frustrating we havent got a look in those craters yet. What is or is not in those polar craters makes such a difference, yet we are talking about committing to HLVs and gateways before the first little robotic mission actually makes it there.

I guess reusability is not too hard if you just mean reuse a few times. I would be more excited by longer stays though, and the longer the stay, again the less call for reuse.

For example, think how the time-on-moon/dollar efficiency apollo could have produced if we had two missions to the same location, one manned, the other with nothing but consumables for the manned mission. Also think how time-on-moon:risk ratio is improved by less times shuttling to and forth. Admittedly this only makes sense now that we have one good location worth visiting repeatedly, and where you can stay longer than a lunar night without boiling or risking being caught in a solar flare.

(im taking this thread too far off topic though.. only want to add the lander-first approach does not negate an L2 argument, but the L2 portion does not necesarily have to be on the critical path either, if we are prepared to accept that landing large payloads long before developing even the ability to land people is not a conspiracy to replace HSF with robots.)

[HappyMartian]

....
(im taking this thread too far off topic though.. only want to add the lander-first approach does not negate an L2 argument, but the L2 portion does not necesarily have to be on the critical path either, if we are prepared to accept that landing large payloads long before developing even the ability to land people is not a conspiracy to replace HSF with robots.)

The Altair descent stage based hydrolox Cargo Lander design gives a potential large reusable hydrolox LM lander that does not need an initial propellant supply on the Lunar surface if it is not heavily loaded with cargo and does not have to do the Lunar orbit insertion burn for the heavy Orion Altair stack.

'All' you would need is propellant supplied by a tanker that came from Earth meet the Cargo LM in Lunar orbit and you would have your ride down to the Lunar surface and back up again on the reusable Cargo LM. Not a bad deal, is it?

Only for large loads going down to the Lunar surface would you then need a supply of propellant on the Lunar surface for the Cargo Lander. However, once the robots get Lunar propellant in production, the 'garage' parked large one way Cargo hydrolox Landers start to get used again.

Anything going to L2 adds risk and is a waste of money and adds nothing to a viable Lunar exploration effort.


See TRAJECTORIES TO/FROM THE EARTH-MOON LAGRANGIAN POINTS L1 AND L2 FOR THE HUMAN EXPLORATION OF THE MOON  By Raśl Cadenas, Carlos Corral van Damme, Simone Centuori     that Proponent attached to his post at:

Cheers!


Edited.

[Solman]

The mass of the concentrator is not small. I used to be a big fan of concentrator systems, but you can get comparable specific powers (or even better) by using efficient thin-film techniques. Concentrators have had problems in the past.


Though they are a pretty good solution for outer solar system missions, since low-light efficiency is disproportionately less, so increasing the intensity helps a lot.

 The concentrator systems used so far have had concentrators with diameters around one foot. The kind L'Garde has demonstrated is 13m in dia. and has a mass of 15kg. I urge you to scan thru the data sheet I referenced - in it you will see depicted a solar installation in Australia that uses such large concentrators with Spectrolab triple junction solar cells which have been operated at nearly 1000 suns as referenced in the same data sheet.
 Your point about concentrator's utility in the outer solar system is well taken - the L'Garde type would deliver around 700W thermal/kg at Jupiter which at 40% efficiency would give 280W electric for a small increase in mass for the PV.
 I had thought you would notice the need for active cooling and argue that that would add considerable mass. I believe that since the L'Garde concentrator structure is formed by over pressurizing aluminum tubes this inflation/rigidizing gas could be circulated thru the structure to radiate heat away. The design would of course have to be tweaked a little for this.
 This is truly game changing wouldn't you agree?

[A_M_Swallow]

Assume we decide to have a cislunar gateway and we stage our missions from there. What is the difference between an expendable and partially reusable lander if they are both staged from the same place, either a Lagrange gateway or one in polar orbit? List all differences.

I am not going to fall into the trap of trying to list all the differences, someone will think of another one.  Here are a few major ones:

Expendable lander are normally one way so can have smaller fuel tanks; a delta-V of 2.52 km/s instead of 5.04 km/s.  (The second generation may be refuelable at both stopping places but not the first generation.)

Expendable landers are lighter than reusable since the can have smaller (= lighter) fuel tanks carrying less propellant.  Their structure can be weaker since they only have to survive one landing, the reduced wet mass may allow a reduction in the number of engines.

Reusable landers need to be equipped with a docking system permitting it to dock with other spacecraft such as the EML spacestation or Orion.

Anything that docks/berths needs GN&C software and avionics that can find the other spacecraft, find the docking port and perform the docking.  An expendable lander could be launched directly from Earth, so it only needs to land.

A reusable spacecraft needs an in-space refuelling system.  This may only be fuel caps that a robot can open but they and pipes to the fuel tanks have to be there.

Every landing needs a new expendable lander to be made and paid for.

The expendable Apollo landers left the descent stages on the Moon's surface.  There are various possibilities but a fully reusable lander design is likely to combine the descent and ascent stages into a single stage.

[A_M_Swallow]

The propellant is not negotiable because it isn't a political issue and making it a political issue could easily destroy this small scale robot based doable Lunar ice ISRU propellant project.
{snip}

Not negotiable is OK with me.

The propellants are methane/LOX for the main engines and hydrogen peroxide for the RCS.

(Happy Martian you did realise that you were the one who was negotiating?)

[A_M_Swallow]

Air startable? 

Im glossing over the restartable detail, but a one way lander seems a good basis for a two way descent/ascent vehicle if fuel can be produced on the surface. I just mean on the moon, without air, descent and ascent are pretty similar I guess? If it can land from orbit, it can be refueled and ascend to orbit? Admittedly you might not have abort to orbit.
{snip}

Why NASA calls vacuum startable engines air startable belongs on a different thread.

The SSME for instance is not air startable so it is unsuitable for use on even the biggest landers.

[Robotbeat]

Huh? Almost every spacecraft since the dawn of the space age has been using hypergolic propellants especially for in-space maneuvering, almost always hydrazine-based. The ISS is essentially a hypergolic propellant depot, with propellant delivered via ATV and Progress.

Ha! That's what I get for posting at 4am while listening to The Space Show. I was sure he said cryogenic.... please disregard my sleep typing.

[go4mars]

I've been trying to read up on silane and related longer-chain hydrosilicons recently (I wish there was more to read about them), and was reminded of Zubrin's note in "The Case for Mars" about a 1050 seconds isp silane and carbon dioxide rocket hopper for Mars.  SiH₄ + 2CO₂ goes to SiO₂ + 2C + 2H₂O.    Which would be 280 seconds isp if the CO₂ wasn't atmospheric. 

What would be the main challenges in creating a CH₄/oxygen engine that could also run on SiH₄? 

I'm thinking in the context of a methane oxygen MCT (or big lander of some kind) arriving at Mars with both tanks full of silane, with 1050 isp for the landing burn, but can be refuelled with either methane and oxygen or just silane. 

Also, a related general question which I haven't seen addressed on this forum - silane also reacts spontaneously with oxygen in Earth's atmosphere.  Why don't rocket first stages use silane and atmospheric oxygen as a matter of common practice on Earth in order to boost isp up near 1000? 
Is it all the nitrogen dulling the reaction?  If so, I would think a centrifuge would be lighter than a giant tank full of oxidizer...   

[metaphor]

I've been trying to read up on silane and related longer-chain hydrosilicons recently (I wish there was more to read about them), and was reminded of Zubrin's note in "The Case for Mars" about a 1050 seconds isp silane and carbon dioxide rocket hopper for Mars.  SiH₄ + 2CO₂ goes to SiO₂ + 2C + 2H₂O.    Which would be 280 seconds isp if the CO₂ wasn't atmospheric. 

What would be the main challenges in creating a CH₄/oxygen engine that could also run on SiH₄? 

I'm thinking in the context of a methane oxygen MCT (or big lander of some kind) arriving at Mars with both tanks full of silane, with 1050 isp for the landing burn, but can be refuelled with either methane and oxygen or just silane. 

Also, a related general question which I haven't seen addressed on this forum - silane also reacts spontaneously with oxygen in Earth's atmosphere.  Why don't rocket first stages use silane and atmospheric oxygen as a matter of common practice on Earth in order to boost isp up near 1000? 
Is it all the nitrogen dulling the reaction?  If so, I would think a centrifuge would be lighter than a giant tank full of oxidizer...   

Rockets generally don't spend much time in the lower atmosphere so it's inefficient to have a first stage with what would basically be jet engines, at least if launching vertically.  The atmosphere thins out much too fast for the air intakes to keep up with the oxygen needs of the engines.

The silane reaction might be useful as a sort of jet engine for Martian airplanes, or anything else that stays in the lower atmosphere.  But I can definitely see some problems with building such an engine.  The exhaust is SiO2, which is basically sand/rock, and C, which is soot.  An engine like that would have major problems with corrosion and debris buildup.

[raketa]

Zubrin has some interesting ideas and is very good at thinking outside the box. But...

Nuclear salt water has never been done. It would have to be developed and tested. Controlling the fission reaction in such a dynamic system would be far from simple. The exhaust products would be so radioactive as to make testing on Earth very difficult and expensive. And politically it's difficult to imagine this being acceptable. Even getting the propellant into space where the system would have to be be used, would be hazardous: imagine what would happen during a launch accident where the payload is many tons of uranium salt solution. I don't see being developed.

The transorbital railroad proposal is essentially a government launch subsidy. I've no idea what the political feasibility of such a scheme would be. Whether it would stimulate the launch market remains to be seen. What effect it might have on international trade agreements is another open question.

I think you're right in saying that Zubrin's ISRU ideas are his best. Some are almost mainstream already.

If you use something like dragon spacecraft with launch abort capabilty, it think it will be possible to bring on the orbit.

[go4mars]

...a related general question which I haven't seen addressed on this forum - silane also reacts spontaneously with oxygen in Earth's atmosphere.  Why don't rocket first stages use silane and atmospheric oxygen as a matter of common practice on Earth in order to boost isp up near 1000? 
Is it all the nitrogen dulling the reaction?  If so, I would think a centrifuge would be lighter than a giant tank full of oxidizer...   

Rockets generally don't spend much time in the lower atmosphere so it's inefficient to have a first stage with what would basically be jet engines, at least if launching vertically.  The atmosphere thins out much too fast for the air intakes to keep up with the oxygen needs of the engines.

I had a conversation recently that answered my question - kind of.  He said that the reason silane and CO2 ramjets work well is because the reaction is far less energetic (more of an intense conflagration reaction).  By contrast, Earth's 20% oxygen atmosphere reacts with the silane, with an exothermic detonation so violent that no one has come up with materials to control and make use of the reaction.  That said, this guy was 90 years old and hasn't kept up with advances in materials sciences in decades.  Perhaps the story has changed materially? 

Almost 3x more capable than on Earth (.38 gravity on Mars).  But I agree with Dr. Zubrin that silane ramjets would be the way to go for hoppers.

Silane ramjets are very intriguing. Two questions:
First, I can see where 1000 s Isp could be a big advantage getting to Mars, but do they have advantages just on Mars' surface at "low" speed?
Second, do you know if there is any project afoot to develop them?

First - not really.  Only at ramming speed (speeds above which ramjets work efficiently but still within the Mars atmosphere).  However a small high pressure starting tank or pump, or supersonic jet could potentially be used to make it work at 0 velocity.
Second - No.  But I figured this thread's title is most fitting to the subject.   

While I'm reminded about it, several people on this forum have expressed a view that Zubrin's 1000+ isp silane ramjets aren't practical/feasible, without providing specific rationale as to why.  For those who hold that view, please consider this a formal invitation and proper thread to outline your thoughts on their limitations (or merits).  Diagrams are very welcome.

[Russel]

Its nice to see relative outsiders have some influence on NASAs thinking and I'd like to see more of that. So whilst I have some respect for Zubrin, and I love some of his ideas, I don't necessarily agree with the obsession that we must do it now, with today's tech.

I'm very much in favour of going to Mars and doing so in an orderly and thorough manner. And to some extent doing it in style with a reasonable amount of margin.

I'll also comment (since I'm from Australia) that I don't necessarily see NASA as being the agency that is primarily responsible for going to Mars (though I'm sure they will have a big part). Yes I know that every space agency has it culture. And the enthusiasts also have their obsessions too. But it would be nice to think that before we go to Mars that we've had time for every interested party to put every idea on the table and come up with an architecture that has been thoroughly thought through.

[JulesVerneATV]

Renowned Mars expert says Trump-Musk axis risks dooming mission

https://www.spacedaily.com/reports/Renowned_Mars_expert_says_Trump-Musk_axis_risks_dooming_mission_999.html