Nuclear rocket

RobertDyck · 2012-01-08 20:43:45

I have an exercise. Design a nuclear rocket for Apollo. For this exercise, assume the same mission as Apollo, and the same 1969-1972 vintage CSM and LM. However, design a new launch vehicle using current state-of-the-art nuclear rockets.

Saturn V had a 3rd state that pushed the CSM and LM into a trajectory toward the Moon (TLI = Trans-Lunar Injection). The Service Module captured the Apollo stack into lunar orbit, and circularized that orbit. That same Service Module pushed the CSM back to Earth, (TEI = Trans-Earth Injection). The 1st and 2nd stages of Saturn V lifted the 3rd stage and Apollo stack into Low Earth Orbit. Let's keep the same mission profile.

Equipment:
NERVA = Nuclear Engine for Rocket Vehicle Applications
NASA developed the NERVA engine for a replacement 3rd stage for Saturn V. This was intended to be the TMI (Trans-Mars Injection) stage for a manned mission to Mars. By 1972 all testing was complete, the only test left would have been in space. It was cancelled on January 5, 1973, by President Nixon; part of the cancellation of Apollo. NERVA was studied again in the late 1980s, study completed 1991; they updated the design but no hardware was built. This one would have had an Isp 925 seconds, thrust of 333.4 kN in vacuum, and engine mass of 8,500 kg.

Timberwind 45
US Air Force developed a series nuclear thermal rocket engines intended for ICBMs. They used a pebble bed reactor with radial flow. That means marble size round fuel elements, arranged in a "doughnut" around the "combustion chamber" of the engine. Liquid hydrogen flowed through the pebbles into the central engine chamber, then down through the exhaust nozzle. I had extremely low reactor mass, but fuel elements melted together. This means it was single use only, not restartable. It was developed, but cancelled in 1992. Nuclear protesters heard about it.
In vacuum: Isp 1000 s, thrust 441.30 kN, engine mass 1,500 kg.

Timberwind 75
Larger engine, same series.
Sea Level: Isp 890 s, thrust 654.600 kN = 66,750 kgf. In vacuum: Isp 1000 s, thrust 735.50 kN. Engine mass 2,500 kg.

Timberwind 250
Largest engine in this series.
Sea Level: Isp 780 s, thrust 1,912.300 kN = 195,000 kgf. In vacuum: Isp 1000 s, thrust 2,451.60 kN. Engine mass 8,300 kg.

Timberwind Titan was a launch vehicle that used a core stage with 3 Timberwind 75 engines, and a pair of Upgraded Solid Rocket Motors (USRM) from a Titan 4B launch vehicle. You don't have to use the same design, it's just a hint.

Fuel tanks for rockets have often used aluminum alloy, but other materials are now available. The External Tank of the Space Shuttle used aluminum-lithium alloy. Carbon fibre epoxy has also been used for liquid hydrogen. It can also be used with liquid oxygen, but since carbon and epoxy burns with oxygen it must have a liner. A fluoropolymer plastic called PolyChloroTriFluoroEthylene (PCTFE) is the most impermeable to oxygen of any polymer that can withstand cryogenic temperature. It used to be sold by 3M under the name Kel-F, but they stopped making it in 1995. Honeywell now makes it under the names Aclar (packaging for pharmaceuticals) and Clarus (for aerospace and defence applications). Aluminized polymer is much more impermeable than pure polymer, so aluminized PCTFE can be considered for LOX tanks. You decide what you want to use.

Now design a TLI stage using a NERVA rocket engine. Is it greatly smaller/lighter than a Saturn IVB stage? Would a Timberwind 45 engine work better? Remember, Timberwind is not restartable.

Next design a launch vehicle to lift that stack into LEO: Apollo CSM, LM, and your TLI stage. Use Timberwind engines for ground launch, you decide which ones. All nuclear thermal, or with SRBs? Or do you want to use small SRBs from an Atlas V?

GW Johnson · 2012-01-09 12:46:29

Why not use the bigger NERVA 3rd stage design, which is restartable (and was back then flight-ready), and push more than one C/SM and a whole slew of landers to the moon, and make several landings at different sites, all in one trip? Isn't that a better return for the launch cost and all the trouble of going there?

Trouble with reviving NERVA or Timberwind or Dumbo or any of them is the lost engineering art as just about all of those guys died or retired. Rocket science ain't all science. It's about 50% art that was never written, just carried in the minds of the practitioners. It's about 40% science, all written down somewhere. And it's about 10% blind dumb luck. And that's in production work. It's worse in development - the art factor is a lot higher. If you lack it, you will think think the blind dumb luck factor is huge, and mostly bad luck.

Been there and done it....

GW

RobertDyck · 2012-01-09 15:36:03

The point is nuclear = much smaller rocket. DON'T use a giant like Saturn V.

As for lost knowledge, just sitting and whining will ensure it's lost. That's why the Mars Society spacesuit group had Dr. Paul Webb himself give a presentation at a spacesuit symposium we hosted within a Mars Society convention. Other movers and shakers in the spacesuit industry were there: representatives from Hamilton Sundstrand, ILC Dover, etc. We heard from all of them their latest efforts in the field, but the reason for organizing the whole thing in the first place was to transfer knowledge from Dr. Webb while he was still alive and his faculties were intact. He just celebrated his 80th birthday.

At another Mars Society convention I met one of the engineers who designed NERVA. I was able to talk to him directly. He was still trying to push his engine for a mission to Mars. But there's no better source than speaking with one of the engineers himself.

GW Johnson · 2012-01-09 18:38:36

I didn't make the meeting where Paul Webb spoke. I would like to have. My dad knew him decades ago as a crew escape expert, before all the pressure suit stuff was known as well as it is today. Dad and I were both aeronautical/aerospace engineers.

I did make the Dallas convention this last summer - gave a paper in the "advanced technology" session on doing a whole slew of landings in one trip to Mars. I got to meet 3 of the original NERVA guys at that meeting, and I bought the book that one of them was selling. It matches my memories of NERVA pretty well.

I quite agree that we really don't need to build another Saturn-5. But if we did, one could resurrect the old NERVA upper stage design for it, and do exactly what I wrote in my previous posting, making a bunch of landings in one trip. The new government heavy-lift launcher design is a new Saturn-5-like vehicle based on retreading shuttle technology. But I seriously doubt any government design will ever be more cost effective than they ever were, which is ineffective.

On the other hand, the new Spacex Falcon-Heavy that is supposed to fly next year is priced at around $1000/lb ($2000/kg) of payload, for 53 metric tons deliverable to LEO. There is not one single reason in the world why a nuclear transfer stage, a couple of big capsules, and a whole slew of modern LEM equivalents cannot be assembled in orbit from 2-5 Falcon-9 launches. One trip, a bunch of landings. Same as Mars.

There is also not one reason in the world why a smaller nuclear upper stage could not be fitted directly to Falcon-Heavy itself. The options are wide open.

GW

i

Rune · 2012-01-10 06:53:59

RobertDyck wrote:

I have an exercise. Design a nuclear rocket for Apollo. For this exercise, assume the same mission as Apollo, and the same 1969-1972 vintage CSM and LM. However, design a new launch vehicle using current state-of-the-art nuclear rockets.

You know you made a loaded question, right? It would make little sense to use a nuclear rocket to substitute just the third stage of a Saturn V. It would make more sense to substitute both the second and third stages with a single nuclear one. Or change the mission scenario to reuse the TLI stage.

I know because one of the most misleading argumentations against nuclear propulsion I've seen (and it was on the wiki) hinged on substituting a S-IVb with a nuclear stage of the same size. So you can see why it got me pissed off.

So, use the tech when the tech is good. Flights with delta-v requirements on the order of 10km/s, in the case of nuclear propulsion. Like Mars and back single stage, Moon and back single stage, single stage mars lander and ascent vehicle, asteroid belt one-way trips...

Rune. I think I'm actually moving the point further along by NOT doing your excercise...

RobertDyck · 2012-01-10 10:37:36

Actually Rune, I argued for what you want. Design the entire launch vehicle as nuclear. But design starts with the destination and works backward. So start with the spacecraft and no launch vehicle at all. Then add the TLI stage. Then and only then design the launch vehicle that can lift that stack to LEO.

GW Johnson · 2012-01-10 11:15:38

What RobertDyck said is exactly correct. Decide what has to go to the moon and land there, and what has to return, first. Then get it there from LEO. Then launch it. That is the correct design sequence. And that's why what you intend to do on the moon so entirely drives the design.

Designing the launch vehicle first (as with NASA SLS, designed by Congressional politics, not engineers) is a wasted exercise.

Getting what you need to LEO need not take one launcher. We now know how to dock things together and assemble very large items in LEO. It's not so much the number of launchers that drives cost, it's cost/mass delivered, and what payload sizes are already flying. Why build a bigger rocket and have to amortize its development costs, if you have a smaller rocket that is "big enough" and already "cheap enough".

Spacex Falcon-Heavy is 53 metric tons to LEO from Canaveral, at roughly $600-1000 /pound (same as roughly $1200-2000/kg). Closest rival is Atlas-5-Heavy at 20-25 tons and more than twice the price/mass.

SLS will be built from retreaded shuttle components by the same entities that built the shuttle, working in the same ways they always did for shuttle. It will never be as cheap as Falcon-Heavy, or even Atlas-5. Shuttle was $1.5B for each 25 ton payload.

GW

Last edited by GW Johnson (2012-01-10 11:16:51)

Rune · 2012-01-10 11:21:58

Well, I kind of got what you wanted to say then, I just thought it was better to just jump to the conclusions and move on from there.

However, maybe I'm a little less turned on by the idea of an all-nuclear solid core launcher. It's not really the radiation, too, although maybe if we used nuclear rockets with the frequency I would like to see rockets used, we could have a problem there. But a nuclear rocket would only make sense to go full SSTO, and then you would run in the usual T/W and throttling problems. I know Timberwind makes it seem as if the T/W problem is surmountable (I think a switch to methane propellant would do wonders, too), but Timberwinds are really complicated things, what with centrifuging the fuel (the nuclear fuel very close to melting temperature) at thousands of RPM's to make the cooling system work, and not being restartables or have a significant design life or throttling capability. Kind of makes the whole reusability thing a bit moot. I'd like more a durable, sturdy design like the old NERVA. Small number of moving parts and active systems (mostly just the turbopump, valves, and the control drums), modest T/W, several restarts, long core life, can even accept several different fuels with no problems.

More sensible, IMO, is to provide a first stage to ease the T/W problem. Some dense liquid or even a solid (though I hate them on reusability grounds) reusable first stage would work better. Plus, it would let you expand the range of places we can launch to directly from the ground, with designs with up to 50% more delta-v in them than current 2-stage vehicles of the same size (and similar mass ratios). You could launch a Dragon-sized capsule, for example, with a Falcon 9-sized rocket, to L1, GEO, or lunar orbit. Or launch more to LEO, or use a smaller rocket, of course.

Upon landing, thrust isn't really an issue, so here's when I appeal to your lack of nuclear fear and ask you to let me land these second stages (almost empty now, so T/W isn't an issue even though the engines run at close to 100% thrust)... "like God and Robert Heinlein intended them to".

Rune. How else?

RobertDyck · 2012-01-10 13:04:16

Well, I got this idea after seeing an old 1950 movie called Destination Moon. They used a nuclear engine, and the whole thing was shaped like a V2. But a scene near the beginning had a retired US Air Force general claim an engineer built a miniature prototype of an engine, the full size one available for this mission would have exhaust velocity of 30,000 feet per second, and thrust of 3 million pounds. That's a movie, but I was impressed how close that is to reality. That exhaust velocity works out to Isp = 923 seconds. If we had that, what would Apollo look like?

This exercise was for entirely expendable rockets. Restart capability for the TLI stage is only required if it's used for LEO circularization, or manoeuvres in transit to the Moon.

Ok, if you don't like this exercise there are a couple more practical ones.

Design a reusable lunar vehicle. Use the International Space Station as a staging location. This makes Earth orbit assembly practical. The reusable vehicle will have to travel from ISS to the lunar surface and back again. Rendezvous with ISS to deliver astronauts and lunar samples. Another spacecraft such as Dragon or Dreamchaser will return astronauts to Earth.

This would be a dedicated in-space vehicle, but there are several safety contingencies. First, in order to keep propellant load down use aerocapture into Earth orbit, then aerobraking to spiral down to LEO. The heat shield does not have to handle re-entry, just enough for aerocapture and aerobraking. I suggest a fabric parasol of Nextel-440. That's the same fabric that the Ames Research Center chose for the latest thermal blanket: DurAFRSI. A parasol held by thin titanium alloy supports, held away from the spacecraft. For safety you would require a small re-entry capsule for astronauts. That's in case aerocapture fails and the craft plunges like Mars Climate Orbiter. But this would be an emergency escape pod, so keep it so tiny it's cramped. A Soyuz descent module would do the job. If you want all American hardware, use the descent module from Northrop Grumman's initial proposal for CEV.

I had looked at using John Wickman's 1980s idea for lunar soil propellant. That used aluminum powder with liquid oxygen; both can be harvested (smelted) from lunar rocks. Two types of feldspar, anorthite and bytownite, can be used as aluminum ore; I gave a presentation about that at the Mars Society convention in Chicago. John Wickman had a couple designs: aluminum powder suspended in LOX. That works and he demonstrated it is not shock sensitive, and combustion does not blow back into the tank. But is mono-propellant safe for a manned mission? Another design of his was bi-propellant: use pressurized nitrogen gas to blow aluminum powder into the engine, LOX comes from a separate tank. This would require bringing some N2 gas from Earth.

I did some rough calculations. I couldn't get any reusable lunar vehicle to work unless it either used lunar ISPP to return to Earth, but that would require lifting fuel to get to the Moon from Earth. To do that you want fuel that has high density, to reduce launch cost. Using different fuels for departure and return isn't practical, not for a reusable ship. However, you could get it to work with a nuclear engine. That would require bringing propellant for return all the way from Earth. Actually, that would be easier, at least for an initial mission, than relying on smelting aluminum from lunar rocks.

The second is for Mars. Design a reusable spacecraft to travel from ISS to Mars orbit and back. Previously I had proposed here on NewMars at mission architecture to do this. Start with an expendable TMI stage, and use the MAV as the expendable TEI stage. Later, once fuel production on a Mars moon is operational, replace them with a reusable TMI/TEI stage. Again use a Nextel-440 parasol for aerocapture in Mars orbit and Earth orbit. Leave the vehicle interplanetary vehicle parked in highly elliptical high Mars orbit, so it doesn't require much thrust to depart Mars orbit. However, at Earth it would aerobrake to LEO in order to rendezvous with ISS. Initially the stack would include a lander that has just a capsule for astronauts plus an all inflatable habitat. If you want a metal wall hab, Ok it's your design. Eventually the landing site will accumulate habs, creating a base. When the mission changes to a reusable TMI/TEI stage, it will bring a reusable Mars shuttle. The best design for Mars would be based on DC-XA.

The surface hab doesn't require a micrometeoroid shield, just protection from dust storms and scuffs from astronauts. Micrometeorites don't survive entry into Mars atmosphere. So use Tenara architectural fabric, which is the exact same material as Orthofabric, the white fabric on the EMU spacesuits used on the Shuttle and ISS, but without the Nomex and Kevlar backing. Nomex is the same stuff as fire fighter's jacket and pants; you don't need a fireproof jacket on a planet with a CO2 atmosphere. Temperature extremes are not as great on Mars: +24°C to -88°C for the absolute extremes, at any location we would land humans. LEO where ISS is parked can go from +120°C to -150°C. Tenara has a twill weave like jeans instead of the double layer plane weave of Orthofabric. That's needed when you don't have the second layer. Again, you don't need the backing on Mars. Tenara is not only lighter than Orothofabric, it's about 1/10th the cost.

The reusable TMI/TEI stage could be all chemical, nuclear thermal, or nuclear electric. Since the entire stage is replaced, there's no commitment to any particular technology.

RobertDyck · 2012-01-10 13:30:43

The big change with Timberwind was a fast reactor design instead of a thermal reactor. As you may be aware, the term “Nuclear Thermal Rocket” refers to the nuclear reactor heating propellant, not the type of reactor. However a “thermal” reactor uses medium speed neutrons called thermal neutrons. A fast reactor uses high speed neutrons. A fast reactor requires highly enriched uranium, while American commercial nuclear reactors use 2% U-235, and Canadian CanDU reactors use non-enriched uranium, sticking with the 0.7% as it comes out of the ground. Technically it's written 235U with the number as a superscript, but BBCode used by this forum doesn't support subscripts or superscripts. Fast reactors require 10% or more, but nuclear rocket engines use >99% anyway. Fast reactors consume their uranium quickly, which allows the reactor to use less uranium. It also “burns” trans-uranic isotopes that are produced when uranium is exposed to neutron radiation in a reactor. And you can use neutron reflector barrels around the reactor. Neutron reflectors not only reduce critical mass required, but these barrels have neutron reflectors on one side, and neutron absorbers on the other. Turning them is all you need to control the reactor. You could stick with axial flow like NERVA, instead of radial flow through a pebble bed like Timberwind. That should fix the agglomeration problem. And reduce the temperature somewhat to ensure fuel elements don't melt. That would reduce Isp from 1,000s to 950s. These changes should result in a reactor much lighter than NERVA, but heavier than Timberwind.

Here's a paper that talks about fast reactors. Note they use hexagonal fuel elements like NERVA. They also mention 950s Isp, that's where I got it from.
http://www.csnr.usra.edu/archives/2006% … actors.ppt

GW Johnson · 2012-01-11 15:22:14

Myself, I would get started going to the moon and Mars with the chemical launchers we have, or soon will have, like Falcon-Heavy. I'd (in parallel) work on resurrecting the old NERVA, since it did everything but actually fly on Saturn-5, and use that type of engine to build single-stage landing craft capable of tail-sitter landings on Mars, and returning to Mars orbit, all in one propellant loadout. Any boat that can do that, can ferry very heavy payloads to and from lunar orbit, down to the lunar surface.

In parallel with all of that, I'd also be working the gas core nuclear thermal rocket ideas, as those could potentially solve the radioactive core problems at performances as far beyond NERVA and Timberwind as they are beyond kerosene-oxygen. I'd probably do the nuclear rocket work on the moon, as doing down here requires no free exhaust, these days. Good reason to go back to the moon, in my opinion. Safe place to do dangerous work of high payoff potential, yet close enough to reach easily with stuff we have right now.

T/W > 10, maybe > 30 at Isp 1500-2500 sec? Good single stage launcher. T/W near 0.1 at Isp near 6000 sec? Good orbit-to-orbit "hot-rod" engine. The bench tests ca. 1969 indicated these things are indeed feasible.

Of course, for really gigantic orbit-to-orbit colony boats, there's good old nuclear pulse propulsion. The bigger the ship, the higher the Isp, and the easier the shock absorber design is. That comes quite a bit later, though.

GW

RobertDyck · 2012-01-11 22:59:03

Gas core nuclear thermal has a lot of potential, but the problem is how do you test it? It spews all its fission fragments with exhaust, creating highly radioactive exhaust. The NERVA guys were worried about radioactive exhaust, and they contained all their fission fragments. Actually NERVA would only have radioactive exhaust if they lost containment, but you test things to see if they work. NERVA did have a problem, the first test had such vibrations from hypersonic flow through the engine that it shook itself apart. Not to pieces, but enough that it didn't work. They fixed it, but that demonstrates the need for testing in any development cycle. So how do you test gas core?

As for nuclear pulse, that means deliberately detonating a nuclear bomb with you inside the blast radius. Very very bad idea. That sounds like suicide. Never mind the ban against nuclear weapons in space. And if a bomb got loose during ground launch, you would drop nuclear bombs who knows where. The government of every country in the world would be convinced you would drop bombs on them, either intentionally or not. Just not going to happen.

But I agree, good old chemical can get us there quickest. I'm the one who asked the Russians in 2000 and 2002 if Energia is available and how much. They did confirm it is available, but didn't confirm the price. But I was told NASA had asked "some time ago" how much; they were told at that time between $60 million and $100 million US dollars to restore infrastructure. NASA has a website that lists Energia at $120 million per launch including the upper stage in 1994 dollars. From that I surmise NASA talked to them in 1994. On April 25 2002 the roof of their vehicle assembly building collapsed, it has never been repaired. I estimate restoring infrastructure today would cost somewhere between $180 million and $200 million. Applying inflation, the per launch cost today would be about $168 million. Notice all this is in millions, not billions. Existing rockets cost a lot less than anything new, no development cost.

GW Johnson · 2012-01-12 15:41:58

How to test gas core?

Short form: same as the solid core program. Put it on a stable thrust somewhere and and fire it. Same as all rockets. The first tests simply cannot be flight tests, that's way too much to bite off all at once.

Yep, the exhaust is radioactive, similar to the Phoebus and Kiwi predecessors to the final NERVA. True enough. A properly-working open-cycle gas core machine will be running at low concentration radioactivity: around 1000:1 hydrogen to uranium fission products by mass. But the total uranium mass fed to the burn will be expelled as fission product mass. The early ones will be much worse, in a concentration sense, until the containment flow scheme works right.

Closed cycle ("nuclear light bulb" designs) will have a clean exhaust, unless the physical containment fails. It will in early testing, occasionally. The problem with closed cycle designs is that the core fission products get retained. I like open cycle better. On shutdown, it's "an empty steel can". Thermally and radiologically, it's "cool" in minutes to hours. Retained cores are dangerous for decades to centuries.

Problem: our rules no longer allow us to free exhaust radioactive plumes. It is possible (in a very expensive facility that we do not currently have) to capture the plume and separate the hydrogen from the radioactive "dirt", and sequester the dirt for disposal.

Wild idea: do it on the moon instead, as free exhaust. Exhaust speeds far exceed lunar escape, and there are no air and water to pollute, or neighbors to bother. It might (!!!) actually be cheaper to do it that way on the moon, instead of plume capture here on Earth.

Even a resurrected NERVA tested here on Earth will have to be tested plume capture.

It's far faster, more effective testing as open plume. The program proceeds much faster and effectively. Resources get concentrated more on the rocket and less on the facility.

Can't do it here? Then do it there! On the moon. It's close enough to reach quickly and with relatively low-performing rocketry. Emergency help is but 3 days away.

BTW, the Th-232 to U-233 breeder cycle, once bootstrapped into operation, yields fission fuel with shorter-lived daughter products. It'll work as a reactor fuel, probably even in nuke rockets, but is not "concentrated" enough to be a bomb. No plutonium in the cycle, either.

I kinda like the concept of a U-233-fed open-cycle gas core engine. It's something easily abortable on launch, and not very dangerous in a crash. At lower power, Isp is near 1500-2500 sec with engine T/W's 10 to 30, or perhaps higher. No waste heat radiator, regenerative cooling is adequate. Higher power, you need the big, heavy radiator: They were going for 6000 sec Isp at engine T/W maybe .05 to 0.1.

I sure wish we'd already done it. The bench tests 40 years ago looked very promising for both gas core reaction controllability, and for open cycle containment by that 1000:1 ratio. That was the target for "perfect containment" at their residence time and burnup rates back then. But it was just some academic-institution bench tests.

GW

Last edited by GW Johnson (2012-01-12 15:45:17)

Rune · 2012-01-12 16:04:37

I took my sweet time replying, but I hope this is worth the wait for you guys. Here are my very rough numbers for a nuclear MTV, as you suggested, for a hypothetical mars mission. Or, as it is single stage and starts and ends its missions at LEO, with 20 restarts on its core before the engine needs replacing, regular martian ferry trips. Enjoy, and don't expect great accuracy, just a rough first-order-of-magnitude sketch.

I have divided the work in, roughly, 6 areas, namely payload (1), engine (2), delta-v and therefore mission profile (3), which gives our mass ratio (4), structural and tankage fractions (5) and then the resulting numbers that come up from all that at the end. Plus, I'll throw out some of the conclusions I get from this grossly unscientific thought experiment at the end. Without further introductions:

1.What's the payload? I'm going to go with a 50ton module. Enough to fit a Transhab, Dragon, and tether system to rotate the whole thing for artificial gravity (with room to spare, Transhab is only 37mT), if it is a crewed flight. Or a decent-sized habitat or lander/ascent vehicle. If you go with the optimist view of things, I think you could use MD's numbers and fit a hab and a lander there, for a certain crew size. Or you can't, and therefore two flights (or two ships) are required to land both a hab and leave a lander on orbit. Also, 50mT is a nice round number, and it fits the lowest cost heavy rocket expected to be in service in the near future. If I cheat a little and look at the end of the post now it's fully written, I can also tell you that if we don't have to bring the payload back (as is the case with surface supplies and landers, and pretty much everything but the hab), the hydrogen-fueled ship can put a little over 115mT ton in mars orbit in this "cargo" version, while the methane-fueled one can deliver over 166mT. In any case, more than enough to deliver both a surface hab and a lander in a single flight, even if the lander is fully fueled.

Note: I will admit it's weird the methane-fueled ship, with it's lower payload fraction, gets a bigger benefit from returning empty, so if someone wants to check my math (did it twice myself), I just worked out two different mass ratios for the two legs of the trip, backwards and without counting the payload on the first one, I'm sure you get the idea. Might just be the mass ratio margin is just bigger for that rocket.

2.Ok, so for the engine I'm going to pick a good old solid NERVA design, with the fuel elements substituted with modern-day CERMET materials and a few overall tweaks bringing the isp to 950, like you said, and maintaining the overall weight and T/W ratio (I think this is conservative enough the resulting engine is reliable as hell). So I get to it, and what's my surprise, DoD did all the hard work fro me. Turns out there is a NERVA gamma, a design study based on the alpha engine that was actually built and tested, with some material upgrades added, including the new fuel elements and an electrical power conversion system. It's also nicely tiny at 3.4mT. The only problem, I am making them run for x4 longer firings. Obviously someone was thinking 0.5G when they designed this engine. Well, I hope they are really well cooled and the core has energy to spare. Here it is:

Engine: 3,410 kg (7,510 lb). Area Ratio: 100. Propellant Formulation: Nuclear/Slush Hydrogen. Restarts: 20.

Status: Study 1972.
Height: 4.46 m (14.64 ft).
Diameter: 1.24 m (4.06 ft).
Thrust: 81.00 kN (18,209 lbf).
Specific impulse: 975 s.
Burn time: 1,200 s.
First Launch: Designed 1972.

The thing eats hydrogen, but the original NERVA was tested with a variety of propellants, including methane. This is important, since cryogenic propellant management technology is in its infancy, while the rest of this work relies on nothing more advanced than what we had in the 70's. Extrapolating the numbers from the old NERVA gives you an isp of 761s in this version of the engine running on methane, and I will keep it in mind as a desirable option. I really don't know what is the lower limit on T/W to call this a "fast burner" that benefits from the Oberth effect significantly (or doesn't incur in excessive penalties), but I reckon at the very least it's got to do significantly more than 0.1G. So for every ~100mT of ship, put a couple of them to have a nice 0.16G acceleration at full weight and plenty of in-built redundancy at <7% weight fraction. Oh, and 20kw in electric power to "run things". Keep in mind that if we make it run on methane, we are blowing considerably heavier gas in our exhaust. If the reactor power was increased, for the same turbopump and related machinery, we would get considerably more thrust, but I won't get into it (even if it makes methane a much less attractive option, T/W-wise).

3.So if we want to get this thing to mars and back, preferably single stage and without needing pesky heatshields so we can aerocapture, what would the delta-v budget look like? From LEO to a Mars capture orbit (big ellipse with great eccentricity that comes within a hair's breath of the atmosphere at the closest point), all propulsively, about 5.2km/s according to wikipedia. You can later, over an extended period of time, slowly aerobrake your way into LMO proper for landing, with perhaps extended stays at selected orbits to study the martian moon's through flybys. Or not, but it increases the lander's reentry speed.

Form LMO to a suitable LEO, 6.12km/s, and you brake propulsively all the way, even though you could just capture propulsively and aerobrake over a few orbits just like you did at Mars, again without any heatshields or extreme precision involved. But I want some margin somewhere, and the astronauts might be in a hurry to get home (and this increases the mission time, at mars you are just cutting on surface time, and there's plenty of that). By the way, I have just pinched numbers from here, if someone wants to work out the real numbers, assume a launch date at least 5 launch opportunities away and work out the actual orbits yourself. But since I used such an approximate approach to everything else, I wouldn't bother.

Incidentally, since the payload requirements I laid out require either two different cargo flights in separate windows for the same ship to be reused (and flight-tested) or building two separate ships (or more, each costs a bit less than the last one), you could go with the expensive safe option and send two manned ships with half the crew each to add more redundancy, and make the lander/ascent vehicle the critical choke point in mission safety (I should work out lander weights one of this days, but it is not that day).

4.Anyhow, adding both up gives you 11,32km/s, which could be a nightmare to fulfill with a single stage chemically, but is a piece of cake four our stupendous nuclear engine. Working the rocket equation, the mass ratio comes up to a shade less than 3.3. Call it 3.5 to have plenty of margin, and fuel to park our ship around a martian moon if the fancy strikes us. That is 2.5kg of propellant for each kg of cargo. If we go with a methane-fueled engine, like I have advocated for a long time, the mass ratio ends up being 4.55, call it 5 to have some margin. 4kg of fuel for each one of cargo, not such a bad trade, considering the fuel is a gazillion times easier to store without no losses, and considerably denser, too, so it would require much lighter tanks. Potential for future ISRU is also greater, but I am sure that thought crossed your mind already.

5.We would ideally like this to be reused a bunch of times, and sturdy as hell to cover any design flaw behind walls of redundancy. So give it a 20% structural fraction. But wait, you say, if you just said the ship has a mass ratio of 5 if methane fueled, that means you just killed that option, right? Well, yes if we built our fuel tanks as sturdy as our main pressure hull. But I am not going to do that. The way I see it, the tanks could be dropped after each mission, be modular and 50mT in weight when fully fueled, and have a structural fraction by themselves of perhaps 10% for the heavily-insulated hydrogen ones. So I'll only count towards structural fraction the unfueled weight of the ship: the payload and the engines, plus the various communication/attitude control/navigation gizmos (a couple of tons at best? I have still budgeted around 6-7mT for them, around 1% of the fully fueled ship, just to be on the safe side and provide both a decent margin and nice round numbers). To put things harder for the methane ship, I'll use the same 10% fraction for the methane tanks, just in case that makes them refuelable at a Phobos station eventually or something. Or good habs to reuse.

So how does it all look like when put together? It is mostly a trial and error process from here, so I'll just give you the results:

With Hydrogen fuel:

Total fueled weight: 525mT
Empty weight: 150mT

Mass breakdown:

Payload: 50mT
Engines (10): 34mT (T/W:~0.16)
Structure: 22.5mT
Comms/Nav/Margin: 6mT
Tankage: 37.5mT
Propellant: 375mT

With Methane fuel (and a relaxed T/W, 'cause maintaining the ratio increases the mass to absurdity):

Total fueled weight: 1.250mT
Empty weight: 250mT

Mass breakdown:

Payload: 50mT
Engines (15): 51mT (T/W:~0.1)
Structure: 42mT
Comms/Nav/Margin: 7mT
Tankage: 100mT
Propellant: 1.000mT

Well, that was that. I have the numpad smoking, that took some iterations to get round numbers (feel free to make up your own)... and correct the stupid mistakes, too. What do I get out of all of this, then? Several things.

First, T/W is important, especially as you get over your exhaust speed in delta-v requirements. The methane case exemplifies it perfectly. The increased mass ratio, though it doesn't seem like much at first glance (3.5 vs 5), ends up costing us a ship two times as heavy with half the acceleration (which will translate in performance losses, and significant ones), because we have to maintain the structural fractions in order for both ships to be as durable and safe. Ok, maybe I over-exemplified it since I used the same weight fraction for both tanks, and a methane tank designed for zero-g could have a structural fraction well below 7%. And the fact that even if it weights four times as much as the other ship, it will be more compact on account of hydrogen being six times less dense.

But you know, the point still stands, this is a mission that is right at the edge of what can be done with <800 isp. On the other hand, this is a mission that is perfectly doable with such an isp, because once you think about it, this ship could be assembled in about 24 flights (20 of them just dumb full fuel tanks to be picked up by the ship), for a launch cost of perhaps ~$2.5B using Falcons. Not bad, not bad at all, considering the R&D cost for the couple of hundred tons of really complicated, aerospace-grade, rest-of-the-ship. Long-term, when we are routinely performing ISRU refueling, this is clearly the way to go at least in the vicinity of mars.

Of course, if we can manage the cryogenic handling of propellants, then oh boy, do we have a ship indeed. Launched in just 11 FH's, refuelable with eight more. A very sustainable launch rate of four FH's each year, plus one every two years for payloads, could maintain a steady stream of missions, each leaving some hardware behind. Add another to loft new engines and spare parts now and then (the engines run out of restarts after 5 round trips at most), and it is still not much. Call it 5 launches a year, for a launch budget of ~$0.5B each year, or about a sixth of the shuttle program. If an engineer with experience on the subject told me the cryogenic problem is solvable within reasonable costs, this is the certainly the way to go in the initial stages of exploration. Later, in the far future, it may very well be that we have become masters at Hydrogen handling and we don't need to switch propellants, or that the physical laws that say "storing liquid hydrogen is difficult" will prove insurmountable and we just switch to something easier to produce and store everywhere (hello ammonia/methane/steam rockets).

Phew, quite a big post in the end. Hope you enjoyed it!

Rune. Or endured it till the end at least... ^^'

Edit: Spotted a couple of mistakes. Didn't change the conclusions on bit.

Last edited by Rune (2012-01-31 08:45:06)

Rune · 2012-01-24 12:44:03

RobertDyck wrote:

The second is for Mars. Design a reusable spacecraft to travel from ISS to Mars orbit and back. Previously I had proposed here on NewMars at mission architecture to do this. Start with an expendable TMI stage, and use the MAV as the expendable TEI stage. Later, once fuel production on a Mars moon is operational, replace them with a reusable TMI/TEI stage. Again use a Nextel-440 parasol for aerocapture in Mars orbit and Earth orbit. Leave the vehicle interplanetary vehicle parked in highly elliptical high Mars orbit, so it doesn't require much thrust to depart Mars orbit. However, at Earth it would aerobrake to LEO in order to rendezvous with ISS. Initially the stack would include a lander that has just a capsule for astronauts plus an all inflatable habitat. If you want a metal wall hab, Ok it's your design. Eventually the landing site will accumulate habs, creating a base. When the mission changes to a reusable TMI/TEI stage, it will bring a reusable Mars shuttle. The best design for Mars would be based on DC-XA.
The surface hab doesn't require a micrometeoroid shield, just protection from dust storms and scuffs from astronauts. Micrometeorites don't survive entry into Mars atmosphere. So use Tenara architectural fabric, which is the exact same material as Orthofabric, the white fabric on the EMU spacesuits used on the Shuttle and ISS, but without the Nomex and Kevlar backing. Nomex is the same stuff as fire fighter's jacket and pants; you don't need a fireproof jacket on a planet with a CO2 atmosphere. Temperature extremes are not as great on Mars: +24°C to -88°C for the absolute extremes, at any location we would land humans. LEO where ISS is parked can go from +120°C to -150°C. Tenara has a twill weave like jeans instead of the double layer plane weave of Orthofabric. That's needed when you don't have the second layer. Again, you don't need the backing on Mars. Tenara is not only lighter than Orothofabric, it's about 1/10th the cost.
The reusable TMI/TEI stage could be all chemical, nuclear thermal, or nuclear electric. Since the entire stage is replaced, there's no commitment to any particular technology.

Done. Only since you designed your own pretty well, I just stole the general architecture idea and reused everything back to LEO, like Von Braun proposed back in the day. Which is, I believe, the final inspiration that got you there, too. BTW, the same "ship" (it's better to think of it as stage, but you sell things the way you have to), different propellant/payload fraction, would get you to LLO and back without problems, so also good for other architectures you mention.

Maybe I'll tackle some of the other parts of the architecture, like the lander/ascent vehicle, or the habs, later. Though designing a hab is just putting a bunch of references together and adding masses, which isn't very fun and has been done to exhaustion.

Rune. Anybody care to poke some holes on it? It's the only way forward! ^^

Last edited by Rune (2012-01-24 12:51:43)

RobertDyck · 2012-01-31 00:29:49

A few points:

NERVA Gamma is great!
For the nuclear engine, use all liquid hydrogen, for both transit to Mars and back to Earth. Assume we have either a water mine on one of the Mars moons, or land on a major permafrost deposit on Mars. The ground penetrating radars on Mars Express and Mars Reconnaissance Orbiter have mapped out those deposits pretty well now.
Use all reusable propellant tanks with the nuclear engine. Only use expendable stages with chemical rockets.
Use aerocapture instead of propulsive orbital capture. Even with a nuclear engine, propellant mass is significant. A Nextel-440 parasol with titanium alloy ribs should mass a lot less. It isn't enough for atmospheric entry, but enough for aerocapture. This would be tested with an unmanned orbiter first.
Did you use carbon fibre composite tanks for liquid hydrogen? They're significantly lighter than aluminum alloy, and even lighter than Weldalite (aluminum lithium alloy). For liquid oxygen carbon fibre composite requires a fluoropolymer liner because both carbon fibre and epoxy burn in oxygen, but that isn't an issue with liquid hydrogen. These tanks were tested/demonstrated with DC-XA.

Last edited by RobertDyck (2012-01-31 00:30:40)

GW Johnson · 2012-01-31 08:44:05

Long-duration LH2 storage can be problematical because of leakage, and because of boil-off. I know gaseous hydrogen will leak right through the steel welding gas bottle over time. Not so bad in liquid phase, but still, we're looking at about 2 years for a round trip, unless one bites the bullet and finds some way to fly much faster.

Here's a really nutty idea: for hydrogen that won't be used until many months later, why not ship it as ice? On-orbit solar thermal to melt, on-orbit solar PV to electrolyze. So it takes a while with a low-powered system. It'll be a while before you need it. It's a very damage-tolerant and minimal-maintenance way to ship your NERVA propellant and your oxygen supply.

I do like the idea of reusable tankage on the nuke ship. Throw nothing away. It can fly many missions, and not just to Mars. Pretty much any design capable of ferrying men from LEO to Mars orbit can be used to visit Venus, Mercury, and some of the NEO's. Refitted with a more powerful engine, it might take men to the main asteroid belt. You need a habitat module big enough for comfy living on 3-year time horizons. I'd suggest a solar flare radiation shelter that is also the flight control deck.

Here's another nutty idea: build the vehicle as a long train of modules. Put the habitat on the end away from the NERVA, and spin it slowly end over end for artificial gravity. That solves a whole host of life support issues associated with zero-gee. 56 m radius at 4 rpm is pretty close to the 1 gee we evolved with.

GW

Rune · 2012-01-31 10:11:44

Hi guys!

I'm just going to put together a bunch of your points and answer them in one go, because I think they all refer to the refueling of the ship, and that eventually leads directly to the overall architecture of the ship's mission:

RobertDyck wrote:

For the nuclear engine, use all liquid hydrogen, for both transit to Mars and back to Earth. Assume we have either a water mine on one of the Mars moons, or land on a major permafrost deposit on Mars. The ground penetrating radars on Mars Express and Mars Reconnaissance Orbiter have mapped out those deposits pretty well now.

and

RobertDyck wrote:

Did you use carbon fibre composite tanks for liquid hydrogen? They're significantly lighter than aluminum alloy, and even lighter than Weldalite (aluminum lithium alloy). For liquid oxygen carbon fibre composite requires a fluoropolymer liner because both carbon fibre and epoxy burn in oxygen, but that isn't an issue with liquid hydrogen. These tanks were tested/demonstrated with DC-XA.

and

RobertDyck wrote:

Use all reusable propellant tanks with the nuclear engine. Only use expendable stages with chemical rockets.

Well, I just budgeted 10% of the fuel's weight to the tanks and left it at that. Should be enough for those recommendations, right? As I said, all the work is approximate, which is the best I can do without actually designing a ship's components. Cryogenic handling over years is also a big source of weight, and even then you can't expect the system to be perfect so you will lose some percentage. And making them reusable may, or may not, be the best idea.
If we are refueling from Earth (and until a refueling stop is built somewhere else, we will), then it could make sense to just leave the tankage section of the fuel tankers docked to the ship and discard the now useless service modules. That way you don't lift any tankage empty. You could even say the tanking is "free", because the refueling ships are going to need them anyway. I'm picturing some tug like progress, Cygnus or ATV, but devoted to fuel. Unless, of course, the unfueled ship is catching dumb tanks launched to orbit, which is risky, because they need a minimum attitude control to be fished, and if you don't circularize, you have very little time to catch them before they reenter. And even then it makes sense that those dumb tanks are also the flight tanks. If our ship is based at the Phobos refueling station and only stops at LEO to pick up cargo and crew and pass maintenance, then that's another thing entirely. But somebody has to build that Phobos refueling station first, and transport each and every piece from earth. By the way, note I don't need no fancy fuel depots. The ship is the fuel depot itself, as it should be.

RobertDyck wrote:

Use aerocapture instead of propulsive orbital capture. Even with a nuclear engine, propellant mass is significant. A Nextel-440 parasol with titanium alloy ribs should mass a lot less. It isn't enough for atmospheric entry, but enough for aerocapture. This would be tested with an unmanned orbiter first.

I think you mean aerobrake, in spirit at the very least. Aerocapture is doing it in one go, and is quite a hairy proposition (not impossible, though). Aerobraking over several passes once you have captured propulsively still saves you most of the capture fuel (capture into high mars orbit is a couple hundred m/s, and I expect on Earth it would be similarly low), and has been done by unshielded satellites. If I didn't use it at the end of the mission is to build in a couple of km/s of delta-v margin, and so I didn't extend the mission for the couple of months required to perform aerobraking maneuvers into a low orbit. Least risk, least complexity, 90% of the benefits: KISS. Also, don't pick your materials until you have a clear idea of the requirements of the component. That one was hammered on me by my teachers.

GW Johnson wrote:

Long-duration LH2 storage can be problematical because of leakage, and because of boil-off. I know gaseous hydrogen will leak right through the steel welding gas bottle over time. Not so bad in liquid phase, but still, we're looking at about 2 years for a round trip, unless one bites the bullet and finds some way to fly much faster.

My point exactly. Which is why I offer methane as an acceptable, storable variant. Ammonia could work too, but has less performance. Plus, everywhere you find water ice, you also find methane, including the lunar poles if we go by LCROSS data. Processing it is waaaay easier. Only fractional distillation required.

GW Johnson wrote:

Here's a really nutty idea: for hydrogen that won't be used until many months later, why not ship it as ice? On-orbit solar thermal to melt, on-orbit solar PV to electrolyze. So it takes a while with a low-powered system. It'll be a while before you need it. It's a very damage-tolerant and minimal-maintenance way to ship your NERVA propellant and your oxygen supply.

Errr... you mean for a chemical H2/LOX rocket, right? Because if not, carrying ice that is only 11% propellant by weight is just nuts. And won't give you a mass ratio close enough to go to mars, ever. You don't need that much oxygen! And the hydrogen will still escape as it is being produced, if you can't produce it quickly enough. More Oxygen you need to carry for nothing. Hell, just run plain water in a NTR and you can do better (isp ~500). But at isp ~500 you either need to refuel on mars or stage the mission and screw reusability.

GW Johnson wrote:

I do like the idea of reusable tankage on the nuke ship. Throw nothing away. It can fly many missions, and not just to Mars. Pretty much any design capable of ferrying men from LEO to Mars orbit can be used to visit Venus, Mercury, and some of the NEO's. Refitted with a more powerful engine, it might take men to the main asteroid belt. You need a habitat module big enough for comfy living on 3-year time horizons. I'd suggest a solar flare radiation shelter that is also the flight control deck.
Here's another nutty idea: build the vehicle as a long train of modules. Put the habitat on the end away from the NERVA, and spin it slowly end over end for artificial gravity. That solves a whole host of life support issues associated with zero-gee. 56 m radius at 4 rpm is pretty close to the 1 gee we evolved with.

Have I already pointed you here? If not, there you can find a Transhab designed for mars missions (so 18 month habitability, and can be extended to 24). Bottom line, as you can see on page 61, the system clocks in at 35mT for a crew of six, and is already designed for 1g at the end of an extensible boom (clever design, BTW) and a power-rich environment. And a solar radiation shelter, EVA systems... the works, of course. The "control deck" can very well be a laptop with wi-fi and LAN access to the rest of the ship, BTW.

Rune. Is glad to talk nukes!

GW Johnson · 2012-01-31 15:22:34

Hi Rune:

I took a look at the NASA report that you sent as a link. The "fire baton" idea is exactly what I was talking about, and I noticed they were using the very same 1 gee and 4 rpm max figures that I use. The transhab looks great except that it is way too small for 6 people -2 years. On a per person basis, Skylab was around 5-6 times the volume in transhab. So, I disagree with their criterion of 25-30 m^3 actually habitable volume per person. It needs to be much closer to 100+ m^3 person to stay sane and fit cooped up like that, the more the better. If one assumes Skylab had half its gross volume habitable, that's 165 m^3 habitable volume for each of the 3 on board. Transhab is just plain too small for a Mars trip.

GW

RobertDyck · 2012-02-02 00:03:32

Rune wrote:

RobertDyck wrote:
Use aerocapture instead of propulsive orbital capture. Even with a nuclear engine, propellant mass is significant. A Nextel-440 parasol with titanium alloy ribs should mass a lot less. It isn't enough for atmospheric entry, but enough for aerocapture. This would be tested with an unmanned orbiter first.
I think you mean aerobrake, in spirit at the very least. Aerocapture is doing it in one go, and is quite a hairy proposition (not impossible, though). Aerobraking over several passes once you have captured propulsively still saves you most of the capture fuel (capture into high mars orbit is a couple hundred m/s, and I expect on Earth it would be similarly low), and has been done by unshielded satellites. If I didn't use it at the end of the mission is to build in a couple of km/s of delta-v margin, and so I didn't extend the mission for the couple of months required to perform aerobraking maneuvers into a low orbit. Least risk, least complexity, 90% of the benefits: KISS. Also, don't pick your materials until you have a clear idea of the requirements of the component. That one was hammered on me by my teachers.

Aerocapture is a specific type of aerobraking. Mars Global Surveyor, Mars Odyssey, and Mars Reconnaissance Orbiter all used a hybrid: propulsive capture into highly elliptical high Mars orbit, then use aerobraking to drop down to low orbit for mapping. That allows your orbiter to map the current Mars atmosphere, determine its exact altitude for precise and gentle aerobraking. That's how they got away without any heat shielding at all. They didn't drop deeply into the atmosphere, just enough to slow the orbiter a little while not producing heat or stress that would require a heat shield. Aerocapture is aerobraking to capture into orbit in the first place. That's more difficult for a couple reasons: you don't have multiple orbits to map the planet's atmosphere, and you have to shed enough speed to capture into orbit. If you don't slow enough, you'll just whip past the planet and go into orbit about the sun. I don't have numbers, but I've been told that is far more intense. I've read about a "balloot", which is an inflated "balloon" that acts as a heat shield for aerocapture. The question was always whether it could handle the stress and heat. My idea was to use a fabric parasol, much lighter than a solid heat shield yet if a small tear happens it just reduces the effectiveness of the heat shield for that tear. A tear in an inflatable would deflate the entire structure. We saw with the thermal blankets on the last few missions of Shuttle that tears do happen. We also saw that a small tear is not a problem. I mentioned Nextel-440 specifically because it's the outer fabric used for DurAFRSI, the latest thermal blankets developed at the Ames Research Center. If you want the fabric to be unspecified at this point, Ok. But I believe aerobraking for orbital capture (aerocapture) of a spacecraft large enough for a human mission will require some sort of heat shield.

Mars Climate Orbiter was the first attempt at aerocapture. I applaud NASA for doing this, however criticize NASA for not trying again. They found the problem: a US measure to metric conversion error. Engineers gave the altitude for aerocapture in US measure, technicians had to convert that to metric to program the orbiter. NASA didn't say the nature of the error, but I noticed before that NASA had published altitude for ISS in miles. But when I checked that with kilometres that I read previously, the numbers didn't match. It turned out NASA was publishing altitude in nautical miles. Apparently their former navy engineers were still using nautical miles. Anyone who's not navy won't even know what a nautical mile is, much less assume nautical miles when given an altitude written as miles. NASA has since fixed that: they ordered everyone to use metric exclusively. With that corrected, they should have sent another orbiter to demonstrate aerocapture. They didn't, NASA has proven too spineless.

Rune wrote:

GW Johnson wrote:
Long-duration LH2 storage can be problematical because of leakage, and because of boil-off. I know gaseous hydrogen will leak right through the steel welding gas bottle over time. Not so bad in liquid phase, but still, we're looking at about 2 years for a round trip, unless one bites the bullet and finds some way to fly much faster.
My point exactly. Which is why I offer methane as an acceptable, storable variant. Ammonia could work too, but has less performance. Plus, everywhere you find water ice, you also find methane, including the lunar poles if we go by LCROSS data. Processing it is waaaay easier. Only fractional distillation required.

Don't bring your propellant for return to Earth, all the way from Earth. Produce propellant for return at Mars, either on Mars itself or one of its moons. So LH2 storage is only depot storage until ready to launch. Once underway, the spacecraft will consume propellant quickly for TEI. So long term storage of LH2 is for depot operation only, not on-board the spacecraft.

GW Johnson · 2012-02-02 15:41:12

RobertDyck:

Is Nextel 440 the replacement for the older Nextel 312? Silicate fire curtain cloth. Your parasol idea is intriguing.

Given a low enough ballistic coefficient, I think it might work for aerocapture, or possibly even re-entry to a landing.

GW

Rune · 2012-02-02 16:04:57

Well, I believe aerocapture can be done, but I really don't think it is such a good idea. If I was going to use it, then the parasol concept is probably the way to go, although retaining the launch shroud could also be another mass-efficient way to go, if it is roomy enough like SLS's. In any case, we are talking about a significant (~10%) percentage of the aerocaptured mass. Which is why I believe that plain propulsive capture and standard unshielded aerobrake passes offer 80% of the benefits, with 20% of the complications.

The only bad part is that it would take you months to get to a circular low orbit from a highly elliptical capture one, but you can either use the time to do other things (remote observation and teleoperation of probes in Mars orbit jumps to mind on that leg of the trip), or just get in the lander and make it reenter at a higher speed form the capture orbit while the ship takes the slow route (Erath orbit, for example, still would be easier than a direct return, so a Dragon/Orion heatshield should take it).

RobertDyck wrote:

Don't bring your propellant for return to Earth, all the way from Earth. Produce propellant for return at Mars, either on Mars itself or one of its moons. So LH2 storage is only depot storage until ready to launch. Once underway, the spacecraft will consume propellant quickly for TEI. So long term storage of LH2 is for depot operation only, not on-board the spacecraft.

That implies a significant infrastructure present already. Who put it there, and how? Do you count its cost towards the cost of the subsequent trips? Plus, this way you are effectively staging your mission in half and you no longer need nuclear engines. Plain chemical ones would do for such moderate delta-v's without problem, and would be both cheaper to develop and to operate. The magic of NTR's is that they can allow mars-an-back (or, eventually, earth-and-back) in a single reusable stage.

Rune. Ditto for the moon, BTW.

RobertDyck · 2012-02-02 17:31:06

My mission plan starts with all chemical, and no infrastructure on Mars. Add mining of Demos or Phobos later, and nuclear engines after those mines have propellant in storage.

I have written elsewhere about mining asteroids. A metal asteroid has gold, silver, platinum, and platinum group metals. By-products include nickel, chromium, cobalt, molybdenum, and aluminum; constituents of an alloy called Inconel 617. You also need a little carbon, but extraction of ferrous metals (iron, nickel, cobalt) is most efficiently done by the Mond process, which requires carbon monoxide gas. That's where the carbon comes from. An effective asteroid mine requires 2 asteroids: one metal, the other carbonaceous. The carbonaceous chondrite asteroid will provide rocket propellant as well as carbon monoxide for mining. The trick is to find a carbonaceous chondrite asteroid close to Earth that still has significant quantities of ice. I read that only those asteroids near the orbit of Mars or farther have ice. Both moons of Mars are captured carbonaceous chondrite asteroids. Because Mars itself has an atmosphere that can be used for aerocapture and aerobraking, the least fuel destination to collect rocket fuel may be the moons of Mars. So fuel to fill a fuel depot in Earth orbit may come via robotic (unmanned) space tankers carrying fuel from one of the moons of Mars. This fuel can be sold to the US military to refuel spy satellites. If that's the case, then return to Earth can be done simply by filling fuel tanks from a depot on the Mars moon itself. Refuelling for the next trip to Mars will be from that depot, filled by robotic tanker from a Mars moon.

Last edited by RobertDyck (2012-02-02 17:36:25)

Rune · 2012-02-08 09:04:01

RobertDyck wrote:

My mission plan starts with all chemical, and no infrastructure on Mars. Add mining of Demos or Phobos later, and nuclear engines after those mines have propellant in storage.

I think it would make sense to start using the nuclear engines to build that infrastructure. As I have shown before, they could be extensively reused and tested to launch all that hardware on the way to mars, and you have the time to build them as you are developing the payloads. Later, they can just switch from being refueled (though that should be "repropelled") on earth orbit from earth-based launchers to being refueled at the depots and stop at earth orbit only for maintenance, refueling of the nuclear cores, and loading payloads.

That way, the total launched mass from the Earth surface would be the absolute minimum, and the only problem is that you have to "develop" the nuclear engine in parallel with the payloads. The engines themselves could be test-fired in high Earth orbit, if you are willing to lose some, and it ends up being cheaper than re-building the test facilities for nuclear engines.

Rune. "Develop" as in "find the guys who built it 30 years ago and ask them how they did it".

GW Johnson · 2012-02-08 10:53:14

I agree with Rune that the sooner we get over our irrationality and apply nuclear propulsion, the sooner things will get affordable, no matter the destination. Don't forget, NERVA-type nuclear is not the best that could be done, it was just the only thing that actually ever got done. Big distinction there.

I disagree that the way to test a nuclear engine is flying somewhere out in space. Nope, chemical or nuclear, you start testing on a stable thrust stand somewhere. If every test has to be a flight test, nothing will ever be done. Because it simply cannot be done that way. Engineering reality.

I suggest we start testing nuclear stuff in a deep crater on the moon. Safe place to do it, and close enough to be reached without nuclear power or gigantic rockets. We can do it with what we already have going. Flying to the moon will prove cheaper than building a facility on Earth that can capture and clean the plume from a nuclear rocket engine that leaks radioactivity. Open plume nuclear tests on Earth are no longer allowed.

GW

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#20 2012-02-02 00:03:32

Re: Nuclear rocket

#21 2012-02-02 15:41:12

Re: Nuclear rocket

#22 2012-02-02 16:04:57

Re: Nuclear rocket

#23 2012-02-02 17:31:06

Re: Nuclear rocket

#24 2012-02-08 09:04:01

Re: Nuclear rocket

#25 2012-02-08 10:53:14

Re: Nuclear rocket

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