Author Topic: NEP AG transit to Mars (Read 58991 times)

SteveMick · « **Reply #60 on:** 12/04/2006 05:56 pm »

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 · « **Reply #61 on:** 12/11/2006 09:38 am »

Quote

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 · « **Reply #62 on:** 12/14/2006 12:07 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

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 · « **Reply #63 on:** 12/14/2006 12:53 pm »

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

neviden · « **Reply #64 on:** 12/14/2006 02:00 pm »

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 · « **Reply #65 on:** 12/14/2006 03:01 pm »

Quote

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 · « **Reply #66 on:** 12/14/2006 04:46 pm »

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 · « **Reply #67 on:** 12/14/2006 09:31 pm »

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 · « **Reply #68 on:** 12/15/2006 07:29 am »

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 · « **Reply #69 on:** 12/18/2006 06:22 pm »

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 · « **Reply #70 on:** 12/18/2006 09:26 pm »

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 · « **Reply #71 on:** 12/19/2006 03:32 pm »

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 · « **Reply #72 on:** 12/19/2006 03:49 pm »

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 · « **Reply #73 on:** 12/19/2006 04:08 pm »

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 · « **Reply #74 on:** 01/04/2007 03:50 am »

....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 · « **Reply #75 on:** 01/04/2007 04:10 am »

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 · « **Reply #76 on:** 01/10/2007 02:05 am »

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 · « **Reply #77 on:** 04/09/2007 02:49 pm »

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 · « **Reply #78 on:** 03/11/2009 01:32 pm »

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 · « **Reply #79 on:** 03/11/2009 07:30 pm »

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.