New methane SC engine.

[krytek]

It seems to enable reusable stages that deeper throttling capability is needed than the current Merlin 1d.

Does methane have any effect on making engines easier to throttle deeper?
And is there an inherent throttling range advantage between GG and SC engines?

No idea how deep, but according to a wonderful Russian document posted by Salo somewhere in the beggining of the thread, there are Russian engines in existence capable of 60-130% thrust.
You read that right, 130%, according to the same paper Methane enables some sort of after burn capability.

[FOXP2]

http://www.scribd.com/doc/26550052/comparative-study-of-kerosene-methane-as-propellants

Has anyone read this yet: it conclude methane ISP increase of only ~10 sec over Kerosene and that the increase engine and tankage mass negates its advantages. The cost of methane is cheaper then RP-1 but that cost advantage is negligible unless significant reusability is achieved. 

[MikeAtkinson]

http://www.scribd.com/doc/26550052/comparative-study-of-kerosene-methane-as-propellants

Has anyone read this yet: it conclude methane ISP increase of only ~10 sec over Kerosene and that the increase engine and tankage mass negates its advantages. The cost of methane is cheaper then RP-1 but that cost advantage is negligible unless significant reusability is achieved. 

It is a study for fly-back boosters. It also compares a RD-180 against a conceptual SE-12 methane engine. Given those parameters it is not too surprising that methane does not offer any advantages.

A key statement in that study seems to be:

In contrast, simulations have shown that a fuel-rich staged combustion cycle is not feasible under realistic assumptions for an engine with such high combustion chamber pressure.

[FOXP2]

I still think the general points of it arguments are valid for spaceX case.

Anyways I agree with the propane alternative, ultrapurified and superchilled to liquid methane temperatures it has nearly the same density as kerosene but would be less prone to coking (though far more then methane which of course can form alkenes), have better cooling properties and might even be cheaper to manufacture then that specialty kerosene that is RP-1. Heck it can even be transported as a liquid under only mild pressure. 

[Joel]

FWIW, I tested running the heavy version (with cross-feed) through the Schilling calculator and then tried to remove the second stage altogether. With 2 stages and cross-feeding booster cores I got 53436 kg to LEO. With 1 stage and cross-feeding booster cores i got 49641 kg to LEO.

All CH4 in both cases and using the numbers suggested by baldusi earilier in the thread (355 s for first stages and boosters, 380 s for second stage). I assumed center core 100 % full when shedding the boosters though.

Anyway, my point is that you might get a decent payload to LEO without any upper stage at all.

[A_M_Swallow]

This comes down to, is lifting the extra 3.795 tonne worth the additional cost of the second stage?

[Joel]

This comes down to, is lifting the extra 3.795 tonne worth the additional cost of the second stage?

Right. Or rather, if you are going to LEO and if you are not going to use up all the 50+ tonnes anyway, why not save yourself the cost of building the second stage?

[go4mars]

Interesting result.  Hard to beat reliability of no staging event (depending on how risky you think dropping the side cores will be).  In a reusability context, a lot of the extra upmass might be used on reusability hardware and extra propellant.

[Joel]

In a reusability context, you save the fuel needed to fly back the center core to the launch site, which must be substantial. And you also do not have the problem of landing with a vacuum optimized US engine.

[simonbp]

But, you need a much, much larger TPS, the mass of which would negate any advantage to the no upper stage option. A heavy upper stage with sufficient TPS would mean a burnout velocity for the core that is only slightly faster than the standard F9.

[Joel]

But, you need a much, much larger TPS, the mass of which would negate any advantage to the no upper stage option. A heavy upper stage with sufficient TPS would mean a burnout velocity for the core that is only slightly faster than the standard F9.

Do you have numbers to back this up? Some back-of-the-envelope calculations: Apollo-CM heat shield was around 1/3 of the landed weight. A modern lightweight material like PICA-X weighs about half of that if I've understood things right. Assume 20 tonnes landed weight for the first stage. That gives you, what, 4 tonnes worth of heat shield? I don't see how that can change the equation.

[go4mars]

In a reusability context, you save the fuel needed to fly back the center core to the launch site, which must be substantial.

True.  I was thinking in terms of re-entry burn, active stage orientation, possibly a plume slip-cushion when it hits atmosphere, and the propulsive landing back at the pad.  Doesn't change your premise, but brings capability back into the EELV range (I would guess).  Which would still be worth doing if it can be done.

[Robotbeat]

But, you need a much, much larger TPS, the mass of which would negate any advantage to the no upper stage option. A heavy upper stage with sufficient TPS would mean a burnout velocity for the core that is only slightly faster than the standard F9.

Do you have numbers to back this up? Some back-of-the-envelope calculations: Apollo-CM heat shield was around 1/3 of the landed weight. A modern lightweight material like PICA-X weighs about half of that if I've understood things right. Assume 20 tonnes landed weight for the first stage. That gives you, what, 4 tonnes worth of heat shield? I don't see how that can change the equation.

A few things: Apollo's heatshield was about twice as thick as it strictly had to be, and it also was returning from the Moon (not just LEO). With modern materials, you could probably cut that number down to just 5% of the landed weight if you were careful.

Also, the first-stage isn't going to be NEARLY that fast at staging. It's just going to be going a few times the speed of sound. I think you're thinking of the UPPER stage, which for the expendable Falcon 9 v1.1 may weigh about 5mT burnout, not 20mT....


Oh, wait, you were talking about using Falcon 9 v1.1 first stage as a SSTO? Horrible idea. It should be possible if you're talking about expendable (though it would have a horrible mass fraction compared to a similar take-off mass TSTO expendable, and it'd be strictly for LEO... it'd be more expensive per kg to LEO than a TSTO), but reusably, it's DEFINITELY not feasible and maybe not even possible.

EDIT: Nevermind, I see you were talking about the center core of the Falcon Heavy. It might be possible to do what you suggest, might even be worth it.

[Joel]

I saw that there was some confusion. The context was the center core of a FH-derived vehicle, using methane as fuel. The extra Isp with methane (I assumed 355 s) lead me to suggest that a second stage isn't necessarily needed for LEO.

[go4mars]

A smallish hydrogen stage on top might be handy in this context.  If they do make a smallish hydrogen stage, it could also act as a third stage when needed (and still fit down highways).