NEP AG transit to Mars

[Chris Bergin]

In works with a powerpoint presentation and animation to download.

We're going to do a short lead in article to bring those people from the news site (still more than the forum - for a strange reason I can't work out) to the forum.

I'll scan some images from the powerpoint I've acquired via the FTP section.

Note this is a March 2003 report, but very interesting reading - noting the devil's advocate approach to the ESAS NTR approach to Mars transit. One for you all to discuss, obviously. Ironically, it has the appearance of the ESAS approach with the landers etc.

[Chris Bergin]

Here's the vehicle from the animation video showing the rotation.

[Chris Bergin]

The report is Earth-Mars Artificial-G NEP Architecture Sun-Earth L2 Architecture (3-Week Parametric Trade Study)

This is an example of one of the pages from the 71 pages in this powerpoint presentation. As mentioned, I'll make it available on the FTP server (I have to given the server loads, and guests taking the link and throwing it on other sites that then use our FTP server externally).

[Chris Bergin]

Here's the link to the download area.
http://forum.nasaspaceflight.com/forums/thread-view.asp?tid=1219&start=1 - which is for presentations and video in L2.

[Mark Max Q]

Seems a little like the SDLV plan, with the CEV being another SDLV. Interesting.

[Chris Bergin]

We're placing another Powerpoint presentation on to the download section - this is from slightly earlier (2002) but has good explantatory background.

58 pages.

[Jason Sole]

Very interesting, and one hell of a read. Might have some comments tomorrow, but impressive to see this was/is/maybe a serious consideration.

[rsp1202]

Graphics indicate masts, etc., are deployable. I hope that means self-deployable or with minimal number of astronaut EVAs for on-orbit assembly, unlike on ISS. There was a shuttle flight or flights that tested deploying such structures.

[David AF]

This is some read and a serious presentation on both reports. Interesting as I think it's useful for seeing how NTR compares?

[nacnud]

I like the picture from 2001, especaly as during that part of the film was supposed to be in zreo g and velcro shoes were simulating gravity.

Not a good start to the report

[Chris Bergin]

Typical, as soon as I use that image on the outline story (which was basically "here's some things mentioned, go read the PowerPoint presentations") you go and say that about the image

[Chris Bergin]

nacnud - 9/1/2006  9:05 PM

I like the picture from 2001,

And I thought you were joking, only to hear it was from the film. I only seem to remember HAL

[Mark Max Q]

Ok, still going over these documents. Firstly what is the reason for moving from NEP to NTR in the ESAS report, compared to where NASA was with the 2002 and 2003 reports?

[HarryM]

Maybe since NERVA went so far along and ground-tested they considered NTR technology more mature.

[Mark Max Q]

HarryM - 9/1/2006  5:09 PM

Maybe since NERVA went so far along and ground-tested they considered NTR technology more mature.

Thanks. Following that up, if given the option with both NEP and NTR, with both along the same level of testing, which is preferable?

[HarryM]

For manned missions the ~12 month NEP transit to Mars (versus the commonly quoted 3 months for NTR) is a big disadvantage. I know which one I'd prefer to be on.  NEP certainly has uses in unmanned deep space, like in the original JIMO concept.

[HarryM]

12 month is what the mentioned PP presentation says (700 days in transit, round trip).

A "fast" Mars mission, Mars Direct with NTR, http://www.astronautix.com/craft/marirect.htm

Mars Direct Nuclear Thermal Mission Summary:

Summary: Low cost; no orbital rendezvous or assembly; dependent upon ISRU propellant production for return; Chemical and NTR options
Propulsion: Nuclear thermal
Braking at Mars: aerodynamic
Mission Type: conjuction
Split or All-Up: split
ISRU: ISRU
Launch Year: 1997
Crew: 4
Mars Surface payload-tonnes: 30
Outbound time-days: 100
Mars Stay Time-days: 550
Return Time-days:  130
Total Mission Time-days: 780
Total Payload Required in Low Earth Orbit-tonnes: 220
Mass per crew-tonnes: 55
Launch Vehicle Payload to LEO-tonnes: 105
Number of Launches Required to Assemble Payload in Low Earth Orbit: 2
Launch Vehicle: Ares

[simonbp]

But the magic is in the thrust; whereas an NTP or chemical manned Mars mission would depart directly from LEO, a practical ion-engine-based mission (solar or nuclear) would have significantly lower thrust, and thus acceleration, and thus spend a significant amount of time climbing up through the Van Allen belts before reaching escape velocity. The spindly NEP spacecraft I've seen also have the distinct disadvantage of being both hard to construct in space (meaning beyond simple docking manoevers) and being to unwieldly for aerobraking (let alone aerocapture) meaning they need a lengthly decceleration burn in order to enter Mars orbit. All this adds up to NEP generally requiring more launches, longer flights, and thus more money.

Could you get around these problems by building a more powerful electric thruster? Yes, that's VASMIR. Would such a rocket plus the nuke to power it end up smaller and more powerful than just an NTR? I doubt it...

Getting back on topic: The trussed, rotating AG would be interesting (for both NEP and NTP) if combined with a direct-entry crew capsule; shortly before reaching Mars, the crew capsule detaches from the truss, gets a distance away from it, and does either a direct atmospheric entry or a series of aerobraking orbits followed by final entry. This might freak out the greens as if the reactor is not on just the right course to skip out on a heliocentric trajectory, it could end up as a radioactive crater...

Simon

[JonClarke]

HarryM - 9/1/2006  7:02 PM

12 month is what the mentioned PP presentation says (700 days in transit, round trip).

A "fast" Mars mission, Mars Direct with NTR, http://www.astronautix.com/craft/marirect.htm

Mars Direct Nuclear Thermal Mission Summary:

Summary: Low cost; no orbital rendezvous or assembly; dependent upon ISRU propellant production for return; Chemical and NTR options
Propulsion: Nuclear thermal
Braking at Mars: aerodynamic
Mission Type: conjuction
Split or All-Up: split
ISRU: ISRU
Launch Year: 1997
Crew: 4
Mars Surface payload-tonnes: 30
Outbound time-days: 100
Mars Stay Time-days: 550
Return Time-days:  130
Total Mission Time-days: 780
Total Payload Required in Low Earth Orbit-tonnes: 220
Mass per crew-tonnes: 55
Launch Vehicle Payload to LEO-tonnes: 105
Number of Launches Required to Assemble Payload in Low Earth Orbit: 2
Launch Vehicle: Ares

Astronautix is an invaluable site but unfortunately the section on Mars missions is full of typos.

The "fast" NTR Mars direct trajectory was 180 days, not 100, the same as the chemical one.  Zubrin used NTR to reduce the LEO mass from 140 to 110 tonnes.

As fars as the NEP proposal goes, it is good to see alternatives to NTR are being considered.  However, the spin gravity idea is completely unnecessary, IMHO.

Jon

[nacnud]

Agreed, for a conjunction type mission with a fast transfer in both directions of 180 days the exposure to micro gravity shouldn't present too much of a problem. This coupled with a long surface stay means you can allow time after landing for adaptation to Martian gravity, if needed.

However for an opposition class mission, where out the out bound journey is still only 180 days there is the short surface stay of 30 days and the long journey home to consider. The short surface stay means that any time needed to recover after a micro gravity transfer has a real impact on the time available for mission objectives. Also the long return leg of 430 days could also be problematic on returning to Earth.

I think the solution is not to use an opposition class mission

Oh and I had to look up the difference between the conjunction and opposition class missions, if anyone is interested this is what I found.

Quick Facts from Red Colony
 

Quick Facts

	Conjunction	Opposition
Outbound Transit Time	180 Days	180 Days
Inbound Transit Time	180 Days	430 Days
Mars Stay Time	500 Days	30 Days
Total Mission Time	910 Days	640 Days
Delta V	6.0 km/s	7.8 km/s
Venus Flyby Needed	No	Yes
Average Radiation Dose	52 Rem	58 Rem
Zero Gravity Exposure	360 Days	610 Days

Diagrams redrawn from International Academy of Astronautics (IAA) Web Site

[MikkelR]

I do not believe that a better alternative than the Mars Direct mission architecture will be possible to come up with. Spending 20-30 days on the martian surface is just not the way to go if you want maximum science output for the cost of the mission.

[Andy L]

Does NTP lose out when placed against NEP, as NEP can also provide electrical power to the ship's systems? Or have I got that wrong on the Electrical subsystem of the NEP?

[nacnud]

That depends on the NTR. There is a vairent called a Bimodal Nuclear Thermal Rocket that can provide electrical power to the spacecraft as well as thrust.

[Andy L]

nacnud - 10/1/2006  9:25 AM

That depends on the NTR. There is a vairent called a Bimodal Nuclear Thermal Rocket that can provide electrical power to the spacecraft as well as thrust.

Thanks, still learning and working that one out with the "Bi" modal.

[nacnud]

Same here, I'm still only halfway through Role of NTR/BNTR/NEP thread

[Avron]

simonbp - 9/1/2006  8:50 PM

 This might freak out the greens as if the reactor is not on just the right course to skip out on a heliocentric trajectory, it could end up as a radioactive crater...

Simon

I have two issue, that someone could provide some insights...

1) How could one safely launch this spaceship, without the huge issues that will be raised by the greens, and its associated political fallout for other nations?
2) What would you do with the ship when the vehicles design life is exceeded, how can it be disposed of? I don't think a dumping it in the Pacific would work, or "Parking" it on a planet or moon would be acceptable?

[Sergi Manstov]

Has Russia or the former Soviet Union worked on such propulsion. Is there any knowledge that could be shared?

[nacnud]

Energia has plans for a solar electric manned Mars mission.

[Avron]

vanilla - 11/1/2006  11:51 PM

Avron - 11/1/2006  9:46 PM
1) How could one safely launch this spaceship, without the huge issues that will be raised by the greens, and its associated political fallout for other nations?

There's nothing to "freak" out about at the launch...the reactor is ice-cold, radioactively.  You don't want it to go critical and begin fissioning fuel if it falls in the drink, but that's not that hard to do, and particularly easy if you use a molten-fluoride reactor, as I mention on the other thread.

Radioactivity is proportional to how fast things are decaying.  Uranium doesn't decay fast (otherwise it would have all decayed away by now).  Fission products decay quickly and are very radioactive.  A reactor doesn't have fission products until you operate it.  You don't operate it until it's in space and in an orbit that will not reenter for thousands of years.

I fully understand the technical angle, but how does one cover the political angle?

[PlanetStorm]

simonbp - 9/1/2006  7:50 PM

But the magic is in the thrust; whereas an NTP or chemical manned Mars mission would depart directly from LEO, a practical ion-engine-based mission (solar or nuclear) would have significantly lower thrust, and thus acceleration, and thus spend a significant amount of time climbing up through the Van Allen belts before reaching escape velocity. The spindly NEP spacecraft I've seen also have the distinct disadvantage of being both hard to construct in space (meaning beyond simple docking manoevers) and being to unwieldly for aerobraking (let alone aerocapture) meaning they need a lengthly decceleration burn in order to enter Mars orbit. All this adds up to NEP generally requiring more launches, longer flights, and thus more money.

Could you get around these problems by building a more powerful electric thruster? Yes, that's VASMIR. Would such a rocket plus the nuke to power it end up smaller and more powerful than just an NTR? I doubt it...

I think NEP has a lot of room for development, way beyond the low-thrust units available today. Have a look at this...

http:/www.scientificamerican.com/article.cfm?chanID=sa006&colID=1&articleID=0004E2EA-0B15-13CC-8B1583414B7F0101

At the moment, this is is just a way of generating intense electron beams for use in laboratories, but what is an intense electron beam to some is a potential propulsion mechanism to others. Perhaps the "spindly" NEP systems we envisage today are just a shadow of what is to come. The key is the nuclear power - with continuous power to spare from a low temperature reactor, there are many options for finding new ways to generate high speed beams, beyond just accelerating particles between two charged plates as in today's EP, or by running the reactor hot as in NTR.

By the way, can anyone help the web challenged amongst us (i.e., me!) who don't know how to paste a link and make it act like one?

[PlanetStorm]

Sorry, the link I just provided is to Scientific American, but you will need a subscription to get the full detail. The basic idea, however, is that you use an intense laser beam to ionize a cloud of helium gas. The rapid heating causes the electrons to seperate from the He nuclei. The electrical attraction between the seperated charges causes an oscillation, with the low-mass electrons moving at high speed relative to the much higher-mass He nuclei. Meanwhile, the laser wave front is continuing to propogate through the helium cloud, and if you get the energetics right, other electrons can be "picked up" and surf the very steep electric wave. The result is a bubble of electrons that follow the propogating laser wave front, and which over small distances (few cm have been demonstrated, but apparently scalable up to metres), the electrons are accelerated to potentially enormous energies.

I don't know if this exact mechanism could ever work as a practical propulsion system for a spacecraft, but it does have the benefit that you are using the lightest possible ejection mass (electrons) accelerated to the highest speeds, so on face value the rocket equation would suggest this is ideal. Either way, it shows that there are mechanisms to heat particles to very high energies without the necessity of very hot thermal reactors or vast (and therefore massive) particle accelerator.

[publiusr]

Those look to be very busy designs--esp. if you had to assemble them ISS style atop cramped EELVs.

Go with HLLV and NTR. It seems a lot more straight-forward.

[Avron]

vanilla - 22/3/2006  11:04 PM

Here on slide 10 of the presentation...they reject continuous thrust vectoring because of the issues of transferring megawatts of power across a rotating joint, and go with the "fire baton" configuration (with body-fixed thrusters), yet that design needs to continuously reorient its thrust vector (which is pointing along the angular momentum vector).  With the Canfield joint, you could keep the vehicle spinning roughly in the orbital plane and then just continually reorient the engines to point along the instantaneous thrust vector (which is relatively fixed in inertial space) while avoiding the issues of transferring power across slip rings.  A simple fat electrical cable will do.

Ok.. go with the  Canfield joint, and "A simple fat electrical cable" ( one that can do whole lot of flexing in temps of space travel - Ref MER.. cables are becomming a problem) ... what do you think is the best design and what stands out?

[dmc6960]

Avron - 23/3/2006  9:56 PM

Ok.. go with the  Canfield joint, and "A simple fat electrical cable" ( one that can do whole lot of flexing in temps of space travel - Ref MER.. cables are becomming a problem) ... what do you think is the best design and what stands out?

If the ship is manned, couldn't a "simple fat electrical cable" be made to be replaceable by an EVA?  Just carry a couple spares around and if one starts giving problems, replace it.

[meiza]

I've done some analysis on electric propulsion for various delta vees. Whether it be nuclear or solar, the ISP can always be optimized, given power density, initial acceleration and delta vee requirements. At non-optimal ISP, it always makes sense (gives lower total mass) to either lower isp and power source mass and add more fuel or up isp and power source and put less fuel. Or fuel=reaction mass, as would be the proper term, if you want to differentiate from the nuclear material.

If I didn't do any mistakes, the equation giving the relation is
deltav=2*k/(a_i*s^2*e^(1/s))
where s = v_ex/deltav, k is power density (W/kg), and a_i is initial acceleration.
Now we see that if you double the power density, you can double the initial acceleration and the other parameters stay the same. Thus a much shorter triptime is reached. (acceleration is always slow with electric propulsion.)

So, what is the power density of those proposed nuclear reactors? Near-term and molten fluoride ones? And what about solar? The Energia (I highly doubt there's some mistakes on the page) study said a 600 ton craft and 15 MW, that gives a rough estimate, if half of the mass is solar cells, 50 W/kg. Though I don't understand what they do if they only get 300N of thrust with that. (Exhaust velocity 100 km/s? Why?)

For example, with a power density of 100 W/kg, and 12km/s of delta v, initial acceleration 1 km/s for every 12 days, a bit over 50% of the craft can be other stuff than power source or reaction mass. Namely, payload. Exhaust velocity is about 40 km/s or the ISP is around 4000. Higher ISP gives worse results, since although the craft needs less fuel, the power source has to be bigger to get same acceleration. These equations don't take into account fuel use that makes the craft lighter later, since it is such a small part of total mass anyway, and this is a rough estimator.

[meiza]

Now we're left with the question, what would be the power density from space nuclear reactors or solar power on a Mars trip. I mean actual numbers. I've been trying to find them on the internet with no luck.

The pdf was vague, it only said 6 megawatts and end mass 89 tons (without fuel). So if 60 tons was reactor and associated equipment, (radiators mostly?) it'd be 100 W/kg. (60 W/kg for actual thrust since the engine is only 60% efficient.) Wouldn't solar cells be in the same ballpark? They are an existing technology and very reliable. Though developing lighter ones would cost some money, it could be useful for other purposes too.

What did they think as reaction mass? Krypton or Argon? Xenon is probably too scarce and expensive for such massive amounts (over 100 t).

[Carl G]

Whoa, old thread, but this is a good one. This would be an amazing ship to build.

[SteveMick]

The numbers you used for solar cell power density are way off. Triple junction PV for use with concentrated sunlight made by the "Solarex" co. currently approach 1KW/kg. The concentrator can easily have a factor of 10KW/kg. and as a result, mass goes up very little as sunlight intensity drops as Mars is approached. Since this tech is developed and at least two orders of magnitude cheaper, I am puzzled as to why NEP would ever be considered for this role.
 Also, a solar electric rocket can operate as a solar thermal rocket to acheive Earth escape from LEO much faster. It really is the best of both worlds and has other advantages besides. The concentrator mirrors can double as communication and/or radar antennas and the intense heat at the concentrator's focus can be used for direct ISRU.
 Please use this as the "competition" for NEP and not a straw man system. I think you'll come to agree that NEP is impractical or at least inferior for Mars transit.
Steve

[meiza]

Hi, I've indeed been in the dark about solar cell mass efficiencies. Are these Solarex cells for space use? References?

[Marcus]

Avron - 11/1/2006  7:29 PM

simonbp - 9/1/2006  8:50 PM

 This might freak out the greens as if the reactor is not on just the right course to skip out on a heliocentric trajectory, it could end up as a radioactive crater...

Simon

I have two issue, that someone could provide some insights...

1) How could one safely launch this spaceship, without the huge issues that will be raised by the greens, and its associated political fallout for other nations?
2) What would you do with the ship when the vehicles design life is exceeded, how can it be disposed of? I don't think a dumping it in the Pacific would work, or "Parking" it on a planet or moon would be acceptable?

1) Fuel launched separately from the reactor packaged in ceramic "indestructible" containment. It still won't satisfy the wackos, but as long as the people with the money are okay with it, we'll suffer the picket signs outside the gates to Canaveral.

2) Keep it away from Earth. The big problem isn't the radioactive fuel. That will probably be in a non-ablative form and difficult to vaporize. It would land in a hot heap somewhere and be easy to clean up. What you don't want to do is accidentally re-enter and vaporize a few cubic meters of radioactive metal structure that's been contaminated by exposure to the reactor flux during the mission. That could measurably raise the ambient level of radiation in the biosphere. It wouldn't kill everyone, but you could probably record a statistical spike in cancer deaths if you did something like that. Say a 0.000167% increase of incidence in the general population or... 10,000 people. Lets not kill 10,000 people.

Crash it into the moon, or at least the radioactive parts. Absolutely safest solution and the lack of an atmosphere makes it attractive in that it won't disperse the contaminated materials. No chance of any measurable amount of radioactive material ever making its way into any biosphere. If the orbits work out better for somewhere else, dump it on the cheapest celestial body with sufficient mass to make sure all the little bits and pieces that break off in the crash don't have escape velocity. Unless you seriously entertain the thought that we might someday have the capability to terraform Mars or Venus yet lack the capacity to deal with a little radioactive contamination.

[TyMoore]

Personally, I've looked at a concept that puts most of the nuclear fuel, or even possibly the reactor on a stripped down verson of a CEV. Use a stripped down capsule that has no lifesupport system, but does have a LES tower and parachute recovery system, ablative heat shield, and of course a recovery beacon. Inside, mount the reactor or nuclear-fuel core in a cask (assuming that part is actually small enough to fit!) inside the capsule. Launch the whole thing on something bigger than the "Stick" so that it can carry enough propellant for a rendezvous with your assembly area. Sure, it's a definate cost, weight, and complexity penalty, but you can use the CaLV to orbit almost all the rest of the vehicle. By using a clever application of a 'boilerplate' CEV for the nuclear fuel transport, you can minimize (almost completely eliminate) any possibility of the fuel core even being damaged in a launch abort situation--let alone resulting in dispersion of radionucleides in the air or water.

I think a dedicated CEV derived "Hazardous Payload Capsule" makes sense.

Of course, this means that whatever reactor is ultimately designed must either be fueled in space by remote, or you've got to send a crew up there to insert a fuel core into the pressure vessel of an NTR or some other power plant. This has never been done before and this would entail a lot of 'firsts' that will admitedly add expense, risk, and complexity to an otherwise expensive, risky, and complex mission....Hmmm...

[kfsorensen]

Space reactors are mandated by federal law to utilize highly-enriched uranium (>93% U-235) as fuel.  The half-life of uranium-235 is hundreds of millions of years, the half-life of the remaining U-238 is 4.5 billion years.  Which means their level of radioactivity is exceedingly small.  In the event of total core vaporization, a couple of hundred kilograms of uranium would be returned to the ocean, to mix with the millions of tonnes of uranium already dissolved there.

You may be confusing reactor fuel (enriched uranium) with the radioactive isotope plutonium-238 that they use in RTGs.  Pu-238 is much more radioactive (with a 78 year half-life) and should not be released to the environment.  The vaporization and dispersion of an entire uranium core (before any fission reactions have taken place) is not a cause for great concern environmentally.

[TyMoore]

I know that before activation of the core, the radionucleide loading is quite small. And personally, I wouldn't lose any sleep over a core burning up in the atmosphere (unless I was working on the project and the loss of a powerplant set my schedule way back!)


As long as you are not trying something really silly like bring back an expended core so that the spent nuclear fuel could be placed in Yucca Mountain (I guess I shouldn't give them any ideas!)  then NTR or NEP are completely safe as far as I am concerned.

Still, it was an idea worth exploring. The effort may either placate or atleast reassure some nuclearphobes who are otherwise supporters of space exploration.

And yes Pu-238 as used in RTGs is a lot 'hotter' (decay activity level wise) than Pu-239. U-235 is actually quite mild compared to Pu-239...

I didn't know about the Federal Law Mandating highly enriched cores--I just assumed that the use of 90-95% Enriched Uranium was more engineering necessity to get the most watts/kg of mass than the Law.

Interesting...

[kfsorensen]

TyMoore - 21/11/2006  4:18 PM

I didn't know about the Federal Law Mandating highly enriched cores--I just assumed that the use of 90-95% Enriched Uranium was more engineering necessity to get the most watts/kg of mass than the Law.

Interesting...

The reason for the HEU mandate (as opposed to launching cores that went critical on Pu-239, LEU, or U-233) is that an HEU core will generate less Pu-239 through neutron absorption than an LEU core.  Pu-239 (and other transuranic) levels in the core need to be kept to a minimum so that in the event that a core experiences a failure as it "climbs out" of the gravity well the orbital decay lifetime of the spacecraft (and its "hot" core) is sufficient to allow nearly all of the fission products in the core to decay to stability.

[Avron]

Marcus - 21/11/2006  3:59 PM

Avron - 11/1/2006  7:29 PM

I have two issue, that someone could provide some insights...

Nearly 10 mths to get a answer.. thank-you Marcus

[SteveMick]

I mistakenly said "Solarex" when I should have said "Spectrolab" - a division of Boeing. They make a variety of cells for space applications and their web address is logically enough www.spectrolab.com
Steve

[James (Lockheed)]

Carl G - 6/11/2006  7:30 PM

Whoa, old thread, but this is a good one. This would be an amazing ship to build.

Sure, but those days of this being possible, due to budgets, are over.

[kfsorensen]

James (Lockheed) - 27/11/2006  6:30 PM

Sure, but those days of this being possible, due to budgets, are over.

If the AG-NEP vehicle comes out to be a lower overall cost than a chemical or a nuclear thermal option, then I don't see why that would have to be the case.  Such an total cost analysis should also include the costs of microgravity adaptation research and countermeasures that would not need to be done if the crew could get the gravity they need for health.

[To The Stars]

vanilla - 27/11/2006  8:09 PM

James (Lockheed) - 27/11/2006  6:30 PM

Sure, but those days of this being possible, due to budgets, are over.

If the AG-NEP vehicle comes out to be a lower overall cost than a chemical or a nuclear thermal option, then I don't see why that would have to be the case.  Such an total cost analysis should also include the costs of microgravity adaptation research and countermeasures that would not need to be done if the crew could get the gravity they need for health.

I'll be a happy man if and when that day comes where this sort of vehicle becomes a reality.

[wingod]

Chris Bergin - 9/1/2006  10:45 AM

We're placing another Powerpoint presentation on to the download section - this is from slightly earlier (2002) but has good explantatory background.

58 pages.

I like this one from Dr. Stuhlinger from 1956

[wingod]

SteveMick - 21/11/2006  2:15 PM

The numbers you used for solar cell power density are way off. Triple junction PV for use with concentrated sunlight made by the "Solarex" co. currently approach 1KW/kg. The concentrator can easily have a factor of 10KW/kg. and as a result, mass goes up very little as sunlight intensity drops as Mars is approached. Since this tech is developed and at least two orders of magnitude cheaper, I am puzzled as to why NEP would ever be considered for this role.
 Also, a solar electric rocket can operate as a solar thermal rocket to acheive Earth escape from LEO much faster. It really is the best of both worlds and has other advantages besides. The concentrator mirrors can double as communication and/or radar antennas and the intense heat at the concentrator's focus can be used for direct ISRU.
 Please use this as the "competition" for NEP and not a straw man system. I think you'll come to agree that NEP is impractical or at least inferior for Mars transit.
Steve

When we put our system together for Langley last year we achieved about 75w/kg for the total system using fairly conservative 33% solar cells.

[SteveMick]

Were you using concentrator type cells from Spectrolab and if so, at what concentration factor?
 What kind of concentrator did you use - one from L'Garde?
 No one argues that traditional non-concentrating type arrays have a mass that is closer to NEP they are so relatively massive.
 Concentrator cells are a real breakthrough and by using the concentrators for solar thermal and antenna duty magnifies this advantage.
 Steve

[kfsorensen]

SteveMick - 29/11/2006  2:47 PM

Concentrator cells are a real breakthrough and by using the concentrators for solar thermal and antenna duty magnifies this advantage.

Using concentrators doesn't do anything about the basic problem of thermal rejection.  If you want to generate 4 MW of power, and you've got 33% efficient cells, assuming everything else is perfect, you need to collect and focus 12 MW of heat energy and then reject 8 MW of it to space.  The radiator will be sized according to the temperature at which you do this rejection, but as a general rule, solar cells don't like to get hot.  Let's assume you have some hot-shot cell that will go to 150 C and still run at 33% efficiency.  Even with a perfectly emissive radiator, you still need to reject 8 MW of heat at something less than 150 C.  But let's be kind and assume you've got a perfectly emissive radiator (e = 1.0) and you've got isothermal heat transfer from the back of your concentrator arrays to your radiator.  You're looking at 6600 square meters of radiator, or a square 80 meters on a side.  If you have to reject at a lower temperature, it gets worse with the fourth power of the temperature.

One of the basic advantages of a nuclear-electric power system is that it is a very dense heat source.  You could couple it to a power conversion system like a potassium-Rankine cycle and reject waste heat at much higher temperatures than solar cells like.

It is also conceivable that you could run a K-Rankine cycle off concentrated solar energy--for typical conversion efficiencies of 20% or so, you'll need 20 MW of input heat to drive the cycle at generate 4 MW of electricity.  That will require ~15000 m^2 of solar collection area (at a flux of 1300 W/m2 and perfect reflectivity) and 430 m2 of radiator area (assuming 16 MW of rejected heat at 900 K emission temperature and perfect emission).

[wingod]

SteveMick - 29/11/2006  2:47 PM

Were you using concentrator type cells from Spectrolab and if so, at what concentration factor?
 What kind of concentrator did you use - one from L'Garde?
 No one argues that traditional non-concentrating type arrays have a mass that is closer to NEP they are so relatively massive.
 Concentrator cells are a real breakthrough and by using the concentrators for solar thermal and antenna duty magnifies this advantage.
 Steve

Sorry I am not a believer in concentrator type systems.  They require precise pointing and the cells degrade faster in a radiation environment.  Also there is the issue of occlusion of the optics in a radiation environment such as going through the Van Allen belts.

I could be wrong but that is my position until someone gives me data otherwise.

[wingod]

vanilla - 29/11/2006  6:51 PM

SteveMick - 29/11/2006  2:47 PM

Concentrator cells are a real breakthrough and by using the concentrators for solar thermal and antenna duty magnifies this advantage.

Using concentrators doesn't do anything about the basic problem of thermal rejection.  If you want to generate 4 MW of power, and you've got 33% efficient cells, assuming everything else is perfect, you need to collect and focus 12 MW of heat energy and then reject 8 MW of it to space.  The radiator will be sized according to the temperature at which you do this rejection, but as a general rule, solar cells don't like to get hot.  Let's assume you have some hot-shot cell that will go to 150 C and still run at 33% efficiency.  Even with a perfectly emissive radiator, you still need to reject 8 MW of heat at something less than 150 C.  But let's be kind and assume you've got a perfectly emissive radiator (e = 1.0) and you've got isothermal heat transfer from the back of your concentrator arrays to your radiator.  You're looking at 6600 square meters of radiator, or a square 80 meters on a side.  If you have to reject at a lower temperature, it gets worse with the fourth power of the temperature.

One of the basic advantages of a nuclear-electric power system is that it is a very dense heat source.  You could couple it to a power conversion system like a potassium-Rankine cycle and reject waste heat at much higher temperatures than solar cells like.

It is also conceivable that you could run a K-Rankine cycle off concentrated solar energy--for typical conversion efficiencies of 20% or so, you'll need 20 MW of input heat to drive the cycle at generate 4 MW of electricity.  That will require ~15000 m^2 of solar collection area (at a flux of 1300 W/m2 and perfect reflectivity) and 430 m2 of radiator area (assuming 16 MW of rejected heat at 900 K emission temperature and perfect emission).

Or you use non concentrating cells and the equlibrium temperature is about 60-70c.

[Avron]

Just wondering what the impact on movement out of the earth-moon system has on both concentrator type system efficiency and the cooling required, while I guess NEP is not impacted.

[kfsorensen]

wingod - 29/11/2006  9:30 PM

Or you use non concentrating cells and the equlibrium temperature is about 60-70c.

Yes but SteveMick was pretty emphatic that the concentrating cells are the ones that should be compared to NEP, and that beyond that even, that concentrating surfaces (inflatable mirrors and so forth) could offer even greater advantages.  I enjoy a good solar dynamic system just as much as the next guy, but I don't think it can credibly compete with NEP for a human Mars mission.

[wingod]

vanilla - 29/11/2006  11:31 PM

wingod - 29/11/2006  9:30 PM

Or you use non concentrating cells and the equlibrium temperature is about 60-70c.

Yes but SteveMick was pretty emphatic that the concentrating cells are the ones that should be compared to NEP, and that beyond that even, that concentrating surfaces (inflatable mirrors and so forth) could offer even greater advantages.  I enjoy a good solar dynamic system just as much as the next guy, but I don't think it can credibly compete with NEP for a human Mars mission.

Sorry I don't buy it.

[SteveMick]

Wow! I thought opinions were supposed to be based on reality and not feelings. If you disagree with the specific power reported by Spectrolab then please talk to them. On the other hand if you agree that the specific power of these cells is in the 1KW/kg range - an order of magnitude improvement over "regular" PV, then their superiority to any NEP system discussed here so far is beyond question.
 The other argument presented here is that current concentrating cells reject heat at a lower temperature and hence require a larger radiator. Since the concentrator support and backing structure is that radiator; there is a certain synergy that mitigates this problem to a great extent. Remember the huge radiators of the JIMO probe were only radiators. The solar radiator can be relatively larger since it doubles as structural support for concentrators that double as antennas and power solar thermal rockets as well.
 I want to emphasize the point that the ability to run in solar thermal mode means much greater acceleration rates and when kinematics favor thrust over Isp can be very advantageous. For LEO to escape a series of thrusts using the thermal rockets at perigee can work very efficiently by taking advantage of the relatively high velocity in this part of the orbit.
 Since the electric prop. sys.'s job in the case of solar thermal/electric propulsion(STEP)is merely to shorten transit time, it remains to be seen what amount of electric power production is optimum, but it need not be anywhere near as large as NEP which ONLY has electric propulsion.
 In addition, there is a high probability that carbon based nanotech cells can operate at far higher efficiencies and temperatures. One design patented in the Eighties used carbon diodes and antennas at nanoscale lengths to convert EM waves of light to elec. just like a radio receiver's antenna does. 60% efficiency or better was claimed in theory. Although no operating temp. was mentioned, you would expect it to be rather high.
 As for the lifetime of concentrator cell, you can carry ten sets of spares before they get as massive as current PV.
 Two to three orders of magnitude cheaper, much more versatile, faster transit times, faster communication, radio astronomy, radar investigations, regolith used as propellent(or just the volatiles) - yeah I guess it can in no way be considered "cedible competition". After all its so superior competition itself between NEP and STEP is not very credible.
Steve

[kfsorensen]

The backside of a concentrating mirror (such as would be used for solar thermal or solar dynamic) is not a suitable radiator, since the concentrator itself is designed to reflect the solar radiation to the receiver at the maximum possible efficiency (reflectivity = 1.0).  Theoretically, none of the heat would then end up at the concentrating surface.

At the receiver, on the other hand, the temperatures will get very high.  For a solar thermal engine, the hydrogen flow is the coolant, but if you desired to run a heat engine like a Brayton, Stirling, or Rankine, you still have to reject the waste heat, which will be between 70-80% of the heat load.  Conducting that heat from the receiver back to the concentrator and rejecting it from the backside will be problematic in many respects.  First of all, there is the sheer quantity of thermal energy to be moved, secondly, a hot back face on the concentrator could distort its optical surface, making its collection efficiency go down.  If the concentrator is an inflatable, then piping heat to it for rejection simply isn't a practical option.

If you use concentraing solar arrays, then the backface of the array can and will serve as the radiator, but you're back to the basic problem of the array area needed and the equilibrium temperature of the array (which in turn alters its efficiency).  To say nothing of pointing and attitude control concerns.

Solar thermal also can't be used to deliver enough of a "kick" at perigee to get you out of the Earth-Moon system at significiant (>2 km2/s2) C3.  Without that capability, much of the astrodynamic advantage of thrusting deep in the gravity well is lost.

[SteveMick]

As you finally came to understand, I was talking about the back of the concentrator - hence my reference to "backing structure.
 As for not being able to thrust at an acceleration rate sufficient to take advantage of perigee kinematics you are incorrect. The thrusts from LEO to the final very elliptical final orbit obviously can take advantage of the kinematiccs even if the acceleration rate is low. You just have more of them. STR's are capable of acceleration rates between 1/100 and 1/10 gee and in the thousand or so seconds near perigee can add as much as 3200 f/s although 1000 or so is more reasonable. For the final thrust it is desirable to provide more acceleration although thanks to electric prop. even a velocity insufficient to reach Mars initially can be overcome. However, even if a chemical stage were used to provide the couple thousand f/sec for a slow trip or more for a faster trip on this last perigee thrust, the advantage is still tremendous. Remember: the perigee thrust need not be anything like the final velocity since that delta V is added to the 24,00mph or so at perigee and the excesss V. is the sqrt of their sum squared. The excess hyperbolic velocity is much greater and delta V's as low as 1500 f/sec at this point are sufficient if Zubrin is correct in "Case for Mars". One option I've been thinking about is adding O2 to the H2 used in the thermal engine on the final perigee thrust to increase thrust or switching to water as STR propellent.
 I also feel that my point about the need for electric propulsion was missed. What I was trying to explain was that the electric power produced does not have to represent 100% of the available concentrated sunlight being used. If the PV is radiator limited(which I doubt) the lower than max. possible power production would have a small effect on transit time.

Steve

[Joffan]

SteveMick - 1/12/2006  10:34 AM

Wow! I thought opinions were supposed to be based on reality and not feelings. If you disagree with the specific power reported by Spectrolab then please talk to them. On the other hand if you agree that the specific power of these cells is in the 1KW/kg range - an order of magnitude improvement over "regular" PV, then their superiority to any NEP system discussed here so far is beyond question.
...
Steve

Steve I had a wander around the Spectrolab site and I couldn't see either this impressive 1kW/kg figure or the earlier 10kW/kg you mentioned... can you direct me to them?

[neviden]

I also couldn't find that figure on spectrolab's site, but i have found some other preposals that mention 1 kw/kg figure, like this one http://www.entechsolar.com/STAIF04.pdf

but i think steve has a point.. you could combine radiator and solar cells, but why not make it like this:

make everything look like a BIG telescope 1 km wide. make a "mirror" from aluminium foil like you would if you would be making a solar sail only curved. focus solar rays onto 40% efficient solar cells. cool those cells (remove 60% that are not converted into electricity) by pumping liquid from cells, through the engine (it does not convert all of the energy to thrust, so it will need to be cooled also), through the mirror suport, to this foil, so that it radiates IR energy into space and then return cooler liquid back to solar cells.

1 km x 1 km x 1,366 W/m2 x 90% reflection = 1,3 MW of solar energy reflected to solar cells
1,3 MW x 40% solar conversion = 500 KW electric power

am I missing something on why this wouldn't work?

[kfsorensen]

The vehicle design in the paper has 4 megawatts of power, or eight times what you've calculated for the solar one.

[neviden]

I have to correct myself.. 1000 m x 1000 m x 1366 W/m2 x 0,9 = 1.366.000.000 W = 1,23 GW of reflected power.. that would produce 500 MW electric power.. you could even skip conversion to electricity and make it solar thermal with hydrogen and isp 900..

Cosmos 1 was a solar sail test craft. It failed to reach orbit, but it had 15 m long wings, 600 m2 of surface and 100 kg weight.. it would concentrate 800 KW of power.. and since it would stay in space it would not have to be very stong and you could build it any size wanted (even 1 km big).. more you could convert to electricity and more efficiently you could cool solar cells, more power you could get.. if you get futher away from sun, then just unfurl more surface and you have the same power..

[meiza]

neviden - 14/12/2006  12:50 PM

I also couldn't find that figure on spectrolab's site, but i have found some other preposals that mention 1 kw/kg figure, like this one http://www.entechsolar.com/STAIF04.pdf

Thank you for the data. This from the article, written in february 2004:
"The state of the art for currently flying space solar photovoltaic arrays is represented by the following key metrics: Areal Power: < 300 W/m2, Specific Power: < 60 W/kg and Operating Voltage: < 200 V."
Those are quite small numbers. Are
The chart about the developments is very interesting too, although maybe it's partly outdated.

What's the current situation with solar cells in december 2006? Any improvements? I have to do some research I guess.

It'd be nice to have a chart with watts per kg. I'd think thin-film cells would do much better there - since sunlight is "free" you can sacrifice some efficiency, just make the cells lighter and you can make them bigger for a net gain. In space of course the problem is supporting the cells (vs on earth where the main issue is cost), but with low accelerations I could imagine they could be very flimsy with little problems if deployment and operations details are taken care of.

[neviden]

I think this is the latest (2006).. >300 W/m2, >300 W/kg, 300-600 V operating voltage, excelent radiation protection http://www.entechsolar.com/SLA-SEP-WCPEC4.pdf

I don't like thin-film cells, because to make them light you make them easily degradable from radiation.. radiation would not make foil less reflective, and since you had few efficient and radiation hardned cells you could reuse "solar tug" for longer (10, 20 years?).. the second thing why i like reflectors is, that you can directly heat hydrogen for high thrust, lower isp manned transfers.. no need for more engines, more power converters, more cooling.. just add bigger mirror and focus on high temperature ceramics pipes filled with hydrogen.. the biggest problem in space is not how to heat stuff.. it's how to cool, since everything is in vacum..

plus.. big solar reflectors/wings can be used to slowly circulize orbit with aerobraking, from HEEO to LEO if built strong enough (like MRO did)..

[meiza]

I'd like to see a Mars SEP design with such cells... You can get pretty high accelerations already with such high power, a few tens of days for trans-Mars (or what's it in greek/latin) insertion? Maybe I could do some calculations. The pointing requirements are quite stiff, one degree or so? But are they only such in one direction? Meaning a simple hinge could do a lot.

If the tech is viable of course... I understand these 1-dim line concentrator arrays (SLA, stretched lens array) haven't yet flown. (Although some of their components have.)

This probably requires starting a new thread, I'll do so if I talk more about this.

[neviden]

the pointing requirement of hubble is a lot stiffer, but we manage to do it.. but you are right, this belongs to new thread

[SteveMick]

With thin film PV, you lose all of the additional advatages I've described from having large concentrator mirrors that can heat propellent or act as antennas. ST/EP is way faster than SEP from LEO for a given vehicle mass.
 This brings up another potentail advatage solar thermal propulsion has. STR's can start at very low initial orbit or even in a suborbital initial trajectory and raise the orbit using hydrogen at 800 sec. Isp or so. The challenge in that case is to design a concentrator that can deploy quickly. Inflation insituform concentrators take a while to harden, but can still function while the hardening is still occuring because the inflation pressure holds their shape untill they do.
 The effect on payload mass to LEO should be dramatic. Perhaps someone more fluent in such calcs could give examples? Acceleration would probably be around 1 f/sec or so I'd guess.
Steve

[TyMoore]

The trouble with trying to use solar thermal in an initially low orbit (here, where the perigee may infact intersect the Earth's surface) and using the Solar Thermal in a circularization burn is solar power (and hence thrust;) the speed of deployment, and of course orientation with respect to the sun. The latter will add an additional constraint to the 'launch window' which is not insurmountable, but increases the chances of technical 'scrubs' from the point of view of synchronizing so many different events.

Also, boosting from a suborbital condition requires more thrust because the duration of the burn is more critical here. The solar concentrating mirror must be deployed and able to track and aquire the sun no more than 1/2 orbit away from second stage seperation, which generally is something less than 25 minutes--it all of course depends upon the actual trajectory used and that is more or less set at the time of Second Stage Engine Cutoff. This may or may not 'push the envelope' of the mirror deployment, aquisition and tracking time. I don't know. But if anything at all goes wrong at deployment (even if it is an otherwise recoverable problem,) then your payload and solar thermal upperstage will definately 'splash' in the ocean! Better to have a better initial deployment and then use a solar thermal stage as a geotransfer stage and apogee kick motor.

The other trouble is depending on the thrust level needed, a commensurate amount of solar energy must be collected. For instance, let's use the figure of 800 seconds of Isp with hydrogen (I think this may be a little bit optimistic,) and a thrust level of let's say 10,000 lbf just for some numbers out of thin air as it were...

First I like to work with metric units: 10,000 lbf=44,480 N. So the mass flow rate in kilograms per second is
mdot=T/(g*Isp) where T = thrust in Newtons. g is acceleration due to gravity at Earth's surface (9.80665 m/s^2,) and Isp is the specific impulse in units of seconds.

Here mdot=5.67 kg/s (or about 12.5 pounds(mass) per second.

The actual mechanical 'power' carried away by the jet of gas is approximately equal to:

Pjet=1/2*mdot*Ve^2 where Ve=the exhaust velocity of the jet, usually approximately equal to Isp*g

In this case with mdot=5.67 kg/s, and Ve=Isp*g where Isp = 800 seconds, and g=9.80665m/s^2, then Ve=7,845 m/s;

so Pjet=1.745*10^8 Watts (or about 234,000 hp!)

Now with typical nozzle efficiencies (better than 60%?, ) heating efficiency of reciever (I'm guessing about 75%) and mirror collection efficiency (95%,) the overall efficiency may be something like: 42.75% (which compared to solar photovoltaics is pretty good!) If 40% is a good overall efficiency, then the mirror must intercept about: 175 MW/0.40 = 438 MW. If insolation at the top of Earth's atmosphere is about 1350 Watts/m^2 this implies that a single mirror must have a collection area of about 324,000 m^2 (or for landscape buffs about 80 acres!) Using two mirrors of 162,000 m^2 each implies that if they are circular then each must have a radius of 227 m (almost 750 ft in radius!) A single large mirror would be about 321 m in radius (2/5 of a mile in diameter!)

This is a very substantial mirror: I doubt seriously that something this size could be deployed as quick as 20 minutes! Reducing our power requirements (and hence thrust) by a factor of ten results in something still very big (still slightly more than 100 m in radius!) but almost useless (1000 pounds-thrust) as an upperstage unless it was used for something like a geosynchronous transfer motor and/or as an apogee kick motor.

If I've done the math right, then this indicates that solar thermal motors, while an interesting concept, are still going to be a bit of a technical challenge to implement.

[SteveMick]

Thanks for that excellent post Ty, but I think you misunderstood somewhat where I was coming from. I mentioned 1 f/sec. acceleration which would imply a vehicle mass of 320,000 lb. for the 10,000 lbf you mention. I was thinking of something considerably smaller as I think a fleet of vehicles should travel together to Mars. Columbus would not have returned if he only had one boat! This means some of the load particularly for surface operations, can be spread over more than one vehicle. Total mass is lessened by the efficiency of ST/EP requiring less propellent vs. chemical which results in such high numbers in Mars transit scenarios and ISRU which can be Phobos regolith. Surface operations can be supported by microwave transmission from orbit to save even more mass vs. some other schemes.
 In addition, even lower acceleration rates still could raise an extremely low initial orbit which should result in considerable gains in payload mass to LEO. I agree that deployment for the suborbital scenario is challenging - so much so that this is the first time I've even mentioned it, but if the concentrators inflate then harden perhaps it can be done.
 Your final point that "solar thermal motors ,,,are still going to be a bit of a technical challenge is baffling  - the deployment issues for the suborbital scenario merit that description, but the low orbit scenario certainly presents no such challenge. Also I am often bemused by the size argument - somehow this is never seen as a problem for solar sails! Concentrator mirrors can be further lightened by nanofabrication of holes smaller than visible light wavelength to make a "light net" with maybe 1/10th the mass. In addition, diamond film simple lens type light net concentrators could have incredibly low mass vs. reflectors. Also size only presents a structural-control problem which I agree is a challenge, but a reasonable one IMO.
Steve

[TyMoore]

Thanks Steve,

Low orbit scenario may suffer from atmospheric drag, as does the ISS. Still, solar thermal motors are interesting. I think Boeing was looking at a solar thermal upper stage, something akin to smaller version of the Centaur vehicle, for placing satellites on a geosynchronous insertion trajectory...

Still, one does get a 'feel' for how much power can be produced by a conventional chemical rocket engine by doing such calculations--this is one reason why I do it. Kind of an "Order of Magnitude" evaluation of the problem before getting into the nitty gritty...

[SteveMick]

I remember seeing an Air force proposal from the late Seventies that proposed a similar idea to the Boeing one you mention.
 Your "doing the numbers" really did put things into the proper perspective and really does show just how powerful the chemical engines are! Thanks again.
Steve

[kfsorensen]

....hopefully getting this thread back on the subject of artificial-gravity NEP vehicles...

It occurs to me that using a Canfield joint for thruster pointing, as I had proposed earlier, would cause the vehicle designer to revisit the thruster placement trade.  If the thrusters each had hemispherical pointing, then it might be far more advantageous to locate them at the reactor end of the vehicle rather than in the middle of the truss.  This would have some beneficial effects.  By locating all the propellant mass down at the reactor end of the truss, some of that propellant might be valuably utilized as both rotational ballast and additional reactor shielding.

Lithium, which is advantageous for lithium-fed MPDs, is an excellent thermal neutron absorber.  The additional mass on the reactor end would enable a longer moment arm for the artificial gravity, and less rotation rate (and Coriolis effect) for the crew.

When mounted at the end of the truss rather than the middle, the thrusters could also be used to help spin up and spin down the vehicle, potentially eliminating the need for additional spinup thrusters down near the habitat module.

[TyMoore]

The only difficulty that I see in placing the thrusters physically near the reactor, is that the reactor will definately produce a 'hard rad' environment near it, i.e., the thrusters may end up getting fried by the relatively intense gamma and neutron bombardment generated by the reactor. Locating the reactor as far away as practical from everything else (like people, computers, landers, and electrical generating equipment) and sticking the propellant between the reactor and 'everything else'  makes sense.


As long as the thrusters are electric, it makes sense to locate them in a more radiation benign environment perhaps between the reactor and the crew compartments. If the vehicle is dual mode with an NTR component, then obviously the thermal thrusters must be located with the reactors. But then again things like pressure vessels and de Levaal nozzles are more robust than the electrical components used in ion or plasma thrusters. Perhaps a combination of both geometries will actually be used...

Also locating electric thrusters away from the reactor also allows for a different set of radiators more 'tuned'  efficiently cool the thrusters and will be much more 'decoupled' from the thermal radiation environment generated by the high temperature radiators used to cool the reactor and power conversion systems.

[Chris Bergin]

I want to link up this forum I've found recently, as it appears to be specific to nuclear propulsion as we are specific to spaceflight:

http://www.energyfromthorium.com/forum/

[kfsorensen]

I think this thread deserves a bump, since the AG-NEP architecture is being discussed in some other threads right now and some of that discussion should probably be here (mostly my fault).

[Archibald]

Hello!

I'm very interested by the AG-NEP proposal, but the download-link in page 1 is dead.

No way of downloading the two Pdfs...

Can someone help ?

[Kaputnik]

Ah I've been looking for this thread!
A few weeks (months?) ago I posted a suggestion for a Mars architecture which relied on NEP, inspired by this very study.

[JohnFornaro]

Here's the link to the download area.
http://forum.nasaspaceflight.com/forums/thread-view.asp?tid=1219&start=1 - which is for presentations and video in L2.

Is there a different level of membership?  when I hit this link, I am told it is not for me to view.  Thanks

[Kaputnik]

Yes, there's the Level 2 area which is subscription only.

[mlorrey]

For manned missions the ~12 month NEP transit to Mars (versus the commonly quoted 3 months for NTR) is a big disadvantage. I know which one I'd prefer to be on.  NEP certainly has uses in unmanned deep space, like in the original JIMO concept.

Shouldn't the transit be faster than 12 months? I thought it was 6 months each way. A Hohmann transfer should leave Earth when it is opposite (aphelion) from the point in Mars orbit where Mars WILL BE when the craft reaches it, which should occur at conjunction.

[JohnFornaro]

Figgered out the subscription thing, and am now subscribed...  Thanks, Chris.

[kfsorensen]

Shouldn't the transit be faster than 12 months? I thought it was 6 months each way. A Hohmann transfer should leave Earth when it is opposite (aphelion) from the point in Mars orbit where Mars WILL BE when the craft reaches it, which should occur at conjunction.

NEP/SEP transfers to Mars have different orbital dynamics than ballistic (high-thrust) trajectories.  Ballistic trajectories leave the departure planet and arrive at the destination planet with significant relative velocities, whereas the EP trajectories leave with little excess velocity and arrive at nearly a matching velocity.  They take longer to fly, but I wouldn't look at this as much of a disadvantage.

[mlorrey]

Shouldn't the transit be faster than 12 months? I thought it was 6 months each way. A Hohmann transfer should leave Earth when it is opposite (aphelion) from the point in Mars orbit where Mars WILL BE when the craft reaches it, which should occur at conjunction.

NEP/SEP transfers to Mars have different orbital dynamics than ballistic (high-thrust) trajectories.  Ballistic trajectories leave the departure planet and arrive at the destination planet with significant relative velocities, whereas the EP trajectories leave with little excess velocity and arrive at nearly a matching velocity.  They take longer to fly, but I wouldn't look at this as much of a disadvantage.

Hmm I'm not sure this is entirely accurate. The studies I saw for VASIMR propelled missions to Mars show it actually takes less time than a Hohmann transfer, with the big time waster being crawling out of Earth's gravity well. After that happens it is more or less heading directly for Mars.
http://www.space.com/businesstechnology/technology/vasimr_rocket_020807-1.html
http://nextbigfuture.com/2007/11/vasimr-engines-plus-200-mw-of-nuclear.html
http://dma.ing.uniroma1.it/users/bruno/Petro.prn.pdf

Constant thrust electric propulsion with their high Isp actually attains much higher average velocities than the chemical or other high thrust/short burn, although the real time savings of such methods is more readily seen in missions to the asteroids and beyond.

Getting nuclear reactors into space is the big stumbling block. I am hoping that Lonnie Johnson's JTEC system could provide a zero moving parts solar thermal power plant thats 60% efficient, which would be a much higher power density than PV or nukes.

[Chris Bergin]

This thread was taken off course, split it. Keep it on NEP AG transit to Mars from now on, or we'll just delete the posts.

[Kaputnik]

Does anybody have a good source of information on these EP Mars transfer trajectories? I'm thinking some sort of table of examples showing the transfer times attainable with a certain thrust and isp. I'd imagine it's a much harder thing to calculate than a simple one-impluse Hohmann transfer.

[kfsorensen]

Does anybody have a good source of information on these EP Mars transfer trajectories? I'm thinking some sort of table of examples showing the transfer times attainable with a certain thrust and isp. I'd imagine it's a much harder thing to calculate than a simple one-impluse Hohmann transfer.

I've done a large number of them.

[Lampyridae]

Does anybody have a good source of information on these EP Mars transfer trajectories? I'm thinking some sort of table of examples showing the transfer times attainable with a certain thrust and isp. I'd imagine it's a much harder thing to calculate than a simple one-impluse Hohmann transfer.

I've done a large number of them.

I guess they'd be a lot tougher than patched conics. Do you use CATS? How detailed are the models?

[GI-Thruster]

Not sure if this is helpful but it's interesting:

http://koti.mbnet.fi/jarmonik/Orbiter.html

[kfsorensen]

A spacecraft like this might be even more suitable for an attempt to visit Ceres. The low gravity of Ceres would play to the strengths of the propulsion technology. I look forward to the information from the Dawn probe on the suitability of Ceres as a future target of exploration.

[Proponent]

Does anybody know what the numerical forward-trajectory model (RKF56) is?

[kfsorensen]

RKF56? Runge-Kutta-Fehlberg 5th/6th order method with error estimation.

[kfsorensen]

Here's a first blush at an animation of the concept.  Fuel tanks and engines aren't shown yet, but I hope you like it anyway:

[kfsorensen]

I did a little digging last week and came to realize that this concept goes back much, much further than I thought.  Ernst Stuhlinger was apparently thinking along the same lines with his work in 1957 and also in 1962:





Here's some great work by Nick Stevens on modern renders of these concepts:

[Robotbeat]

Almost 8 years later! Heck of a necro-bump (70 years if you count the original idea you dug up!). I like it it, Sorensen!