There are lots of reasons to look at methane. Lots of ways it can be used too:
either as a fuel to be burned with an oxidiser, or in a tripropellant rocket, where methane is used for brief high thrust periods, before it switches over to straight H2 for the higher isp. (I gather the Soviets/Russians have looked into tripropellant a lot for spaceplanes like the system 49 and MAKS, for the reasons stated here, lower volume for the same mass, though I think they mostly looked at it for ground-LEO, and used RP. Various wrinkles like this, along with zero/half stages and such are a good way to cheat your way to an erzatz SSTO spaceplane. CH4 would be preferable to RP for the cleaner burn, IMO.)
CH4/H2/O2 tripropellant would work well for a chemical rocket Mars global/orbital hopper.
Another option for an NTR using CH4 is to inject LOX into the exhaust for brief high impulse burns. This is always an option for the "NIMF" described by Zubrin, for self-refueling while grounded at an exploration site on Mars, though in "the Case for Mars" book, he described the attraction of the extra simplicity of using either CO2 (ultimate simplicity on Mars) or CO, with O2 injected for the "afterburner" effect.

The NASA Langley HL-42 crew spaceplane proposal was designed to use CH4/LOX for its onboard OMS and RCS fuels. Reportedly, they were under a "no hypergolics" ground rule. We ccan only assume that they used CH4 instead of H2 for the above stated handling reasons, but I wonder if here was any interest at looking at using RP instead of a cryoogenic...
www.astronautix.com/craft/hl42.htm


-- Edited by john fraz at 23:47, 2007-12-13

Don't agree. Hydrogen is a very efficient fuel, so there's the issue of cryogenics we would just have to deal with the issue. Unless I was forced to use methane because of extraterrestrial supply problems -  I would always favor LH and LOX without which you could not get a space program off the ground. 

Certainly LH₂ was a breakthrough and was necessary for the Apollo missions.  Otherwise we would have had a rocket that would have been impossibly large. I have given the reasons why NTRs should generally be LH₂ based on the relationship between molecular weight and specific impulse.  If you are getting the methane on-site thats a whole different thing. 

I still see the CEV/Orion as an intermediate step (the one step backwards) that we are forced to make. When we finally get back to a reusable shuttle system one factor that must be considered is (as Googlenaut agrees) is density.  When you compare liquid methane to kerosene there is a small but minimal advantage for the methane. That is why it has been the general practice to just use the kerosene.  Based on some discussion that Ive had recently, it turns out that hypersonic scram-jet conceptual designs have avoided kerosene fuel do to concerns about carbonization of the burners.  These concepts have therefore favored methane for cleaner burning or LH₂.

The only way I see possible for a SSTO shuttle that might be developed from this point is some sort of combined cycle air-breather.  Given that the small X-43 is the only flying example for this sort of propulsion system we are some ways away from it.  We need a significant amount of propulsion research before the will be possible. The other big problem is that such vehicles conduct most of their acceleration in the 100-200 K feet altitudes.  There will be considerable drag and heating associated with moving to orbital speeds in this region. Two things that could help are to use the fuel to cool the skins before burning it which could keep the thing from melting and increase specific impulse at the same time and to partially supplement the combustion with onboard LOX. 

If the above is too high risk or upon detailed study turns out to be infeasible, the fall back is a two-stage vehicle with rocket propulsion.  I think that there is a good case for basing this system on a kerosene-LOX propulsion.  To make this development cost-effective, it is key to improve the turn around times and increase sortie rates above the existing space shuttle. The LH₂ has place to much strain on the fuel pumps on the shuttle requiring extensive work between flights.  The less stressful kerosene-LOX should avoid this problem.  Also, with kerosene vs. very cryogenic LH₂ one could use more of the enclosed volume of the winged vehicles for fuel while the LOX would be in one central tank in the fuselage.   We could get Isp of 300 at launch and up to 360 at the end of the burn of the orbiter stage. This should be good enough to achieve orbit with practical propellant factions (rocket version of aircraft fuel fraction).   The higher density of kerosene would compensate by reducing the wetted area and weight of the TPS. 

-- Edited by John at 14:46, 2007-12-20

Wow! Alright--I remember reading about the NIMF--Nuclear Indigineous Martian Fueled rocket--and I thought it was particularly ingeneous in light of the fact that it was designed to use liquid CO2 for reaction mass. The reduced surface gravity and atmospheric pressure lends itself readily to such an application--and liquid carbon dioxide has a pretty high bulk density (about 1.03 times water)--it seems to me that I remember doing some calculations that suggested that NTR with supercritical CO2 ought to give about 200 seconds of specific impulse at reasonable 250-500 psi chamber pressures (I don't have my data sheets handy, but I think I can rework those!) Not 'stellar' performance, but entirely adequate for the application of Martian SSTO. If you put external drop tanks on that thing--you might get enough delta-V for a flight all the way back to earth.

The first page of your paper is located here:
http://pdf.aiaa.org/preview/1991/PV1991_2336.pdf

I don't think I ever read the original paper--should make for interesting reading. I am a fan of the NIMF--a nice balance between simplicity and brute force (provided by the nuclear power source.)
The paper mentions a pebble bed reactor--what was the composition of the pebbles that you were going to use?

I am a fan of using tugsten with tungsten-zircalloy (zirconium alloyed with 93% enriched U-235 to dilute it a bit,) with an all metal core you get higher heat transfer rates and better structural characteristics. This works really well with hydrogen, but liquid CO2 tends to be a bit 'corrosive.'

Yes,
The novelty is being able to use CH4 and CO2 in the same engine.  My design is a hybrid form of Zubrin's NIMF.

The name of the engine under study is called LANTR which stands for Lox Augmented Nuclear Thermal Rocket--which basically uses a nuclear core to heat hydrogen from liquid to high temperature gas, and then expands in part way through a nozzle. Just aft of that primary regeneratively cooled nozzle is a second regeneratively cooled nozzle extension. Where the two meet is where the liquid oxygen is sprayed in. The oxygen 'afterburns' with the hot hydrogen and provides additional mass at velocity to provide thrust augmentation...

Usually when looking at the numbers it looks something like this:

Thrust without augmentation: 15,000 lbf
Isp with hydrogen only: 900 sec
Reactor Power: about 1000 MWt

Thrust with Augmentation: 75,000 lbf
Isp with augmentation: about 500 sec
Reactor Power: about 1000 MWt

These are approximate numbers from memory--but they should be pretty close.

Copies of the the original white paper can be had from:
http://www.engineeringatboeing.com/dataresources/AIAA-2004-3863.pdf

The Pratt and Whitney TRITON design is an exciting concept and deserves further development.

As far as multi propellant flows through the same engine---I'm not sure about corrosion issues--surfaces designed to interact with hot hydrogen and methane may be corroded by superheated steam...perhaps coating the interior gas contact spaces of a tungsten core reactor with an alloy of Platinum-rhodium-rhenium would work. This might not be a bad idea to keep carbon residues to a minimum when processing methane...I'll have to do a little more research on that...

One thing I am fairly sure about--you can't use the same turbopump to handle different flows. A pump designed to move liquid hydrogen will not work for pumping water--fluid mechanics will be too different. A pump that can handle liquid methane ought to be able to pump water, but not vice versa. I suspect that you would need three sets of pumps for three different propellant flows. And God help you if you try to run liquid hydrogen again after running with water--water ice in the cooling jacket and turbopump impellars, not to mention in the bearings would really suck--gives me a cold shudder just thinking about it...


-- Edited by GoogleNaut at 13:53, 2007-12-21

Yeah, you're right. Running some numbers directly from the paper, here's the break down:

For an RL-10 class NTR, directly from the first page of the paper, we get:

Straight NTR-mode:

F=66.7kN (15,000 lbf) thrust,
Isp=911 sec,
which implies a liquid hydrogen consumption rate of about:

m_dot = 66,700 N/ (911 s * 9.80665 m/s^2)
          = 7.5 kg/s

The mechanical jet power (and also very nearly reactor thermal power) is thus about:

Pjet = 1/2*7.5 kg/s * (911 s * 9.80665 m/s^2)^2
       = 300 MW. This will be our 'initial jet power.'

Now if we assume that the increase of jet power due to thrust augmentation is entirely the result of the combustion of available hydrogen we now come to the second part:

From the paper, it is mentioned that when oxygen is introduced into the hot hydrogen stream at a (by mass) mixture ratio of 3 to 4 the thrust increases by a factor of 200%, for a net total of 300% rated thrust, or about:

F augmented = 200 kN (45,000 lbf)

Let's assume an O/F ratio of 4, and that combustion is complete and perfect, and 100% efficient at converting thermal energy into kinetic energy of exhaust products. In such an ideal situation, we have an equation describing the change in kinetic power of the exhaust jet as the gasses move through the combustion zone inside the nozzle:

delta-P = Pjet, final - Pjet, intitial = available combustion power. So solving for Pjet,final we get:

Pjet,final = Pjet, initial + delta P.

Let's get the combustion power next:

From wikipedia.org, we see that Hydrogen has a low enthalpy of combustion (noncondensing products) of -286 KJ/mole (in chemistry the significance of the negative sign indiciates heat is lost or given up in the reaction,) or about 143 MJ/kg.

Now with an O/F=4, and 7.5 kg/s of hydrogen, we need atleast: 30.0 kg/s of liquid oxygen flow. This 30.0 kg/s of liquid oxygen can completely oxidize 3.75 kg/s of the hydrogen, resulting in a combustion power of about:

P combustion=3.75 kg/s * 143 MJ/kg
                   =536 MW.

Thus our total, augmented jet power is about:

P jet,total = P initial + P combustion
               = 300 MW + 536 MW
               = 836 MW.

Now estimating the total augmented Isp from jet power can be easily found from realizing:

Pjet = 1/2 *m_dot *(v_jet)^2 and solving for v_jet, we get:

v_jet = sqrt(2*P jet/m_dot.)

Substituting Pjet=836 MW; and m_dot = 37.5 kg/s (total mass flow rate of oxygen plus hydrogen,) we get:

v_jet = 6680 m/s which corresponds to an augmented Isp = 680 sec of specific impulse--which I consider pretty good. The actual number would in reality probably be something like 5% less (due to combustion losses, boundry shocks, turbulance, etc.) but that still is a pretty good Isp = 647 sec.

It seems to me that if you go to the trouble of using NTR for propulsion, then using a thrust augmentation scheme with aftercombustion with oxygen injection might make sense for departure/capture trajectories, especially with a manned vehicle making a minimum time transit of the Van Allen radiation belts...
Of course only a detailed analysis of the mass cost/benefit from a range of possible mission trajectories will give the final answer, but still, to me at first it looks good...

Ty Moore
Eureka, CA

This has been discussed on other threads--again the problem with any SSTO is thrust to weight ratio--if you get terrific performance but your engine is too heavy, then it won't fly. In a SSTO system the mass margins are very tight, especially for chemical systems, and most especially for SSTO from Earth surface. NTR, even augmented NTR gives good performance, but the thrust to weight ratio compared to high performance chemical engines (such as the U.S. SSME or Russian RD-170 series) is pretty marginal. Also, I'm concerned about scaling--rocket engines are extraordinarily finicky beasts; so are nuclear reactors--they do not scale linearly. Each design must be complete, self contained and self consistant--this is why bigger engines tend not to resemble smaller engines even though they may have the same propellant and pressure heritage...it is just the way the physics and engineering works out...

Bigger reactors generating ten times more power may infact be less than twice as big...

SSTO from Earth's surface may be possible, but it is really, really hard.

I'll take a more detailed look at it--but my first impression is that it won't give sufficient thrust-to-weight ratio to make it work. In the past I have tried to design Reusable SSTO vehicles using both hydrogen/oxygen, and LOX/RP-1 and some that used both--and the the GLOW (Gross Lift Off Weight) to deliver useable payloads tend to get rediculously high (I remember one 'design' that called for a vehicle that massed over 27,000 metric tons at takeoff! The Main liquid oxygen feed line was over 15 ft in diameter!) I haven't tried with LANTR style engine yet...so I'll take a crack at it!