New methane SC engine.

[go4mars]

2e) Long term storage in space for prop depots, or re-start after the long voyage between planets.

[RedLineTrain]

I imagine that having assembled a top-notch engine development team, Musk would not want it to go idle.  After the Merlin 1D and the Super Dracos, what engine developments do they have left?

Cross-feeding the Falcon Heavy. It's a non-trival problem, which is why it's never been done before (unless you count the original Atlas, which had a torturous and expensive development).

Would that require a lot of work from the engine development team?

[ArbitraryConstant]

SpaceX has developed three engines almost from scratch (Merlin, Kestrel, and Draco); they know how run an engineering project. The new SC engine is just not a priority at the moment, and won't be until both a manned Dragon and Falcon Heavy have flown.

I doubt it, just for business reasons.

If they've got Merlin mostly where they want it, and the Super Draco design is pretty far along its development path, they'd have an under-utilized engineering team, probably one of the best in the world. Keeping them idle won't launch Dragon or FH any faster, and SpaceX has the cashflow to keep them busy. So, use them or lose them.

And what would the result be? An inexpensive, high thrust, high ISP engine could be one of the biggest competitive advantages and revenue streams SpaceX could have going forward, the RD-180 of the next few decades. Even if SpaceX fails completely with their own vehicles they could well have other customers for the engine.

[Lars_J]

I imagine that having assembled a top-notch engine development team, Musk would not want it to go idle.  After the Merlin 1D and the Super Dracos, what engine developments do they have left?

Cross-feeding the Falcon Heavy. It's a non-trival problem, which is why it's never been done before (unless you count the original Atlas, which had a torturous and expensive development).

I think the 30yr Shuttle experience should count for something here... Engines being fed from an external (& detachable) propellant source.

[Jorge]

I imagine that having assembled a top-notch engine development team, Musk would not want it to go idle.  After the Merlin 1D and the Super Dracos, what engine developments do they have left?

Cross-feeding the Falcon Heavy. It's a non-trival problem, which is why it's never been done before (unless you count the original Atlas, which had a torturous and expensive development).

I think the 30yr Shuttle experience should count for something here... Engines being fed from an external (& detachable) propellant source.

Shuttle ET sep occurred in vacuum, at nearly 100 km altitude, and after MECO so the shuttle was no longer accelerating. If SpaceX intends to stage FH at a similar dynamic pressure, with the core stage completely cut off, then the shuttle experience can count. On the other hand, if FH is going to stage at a lower altitude (higher Q) and/or with the core still firing, then it's a completely different problem and the shuttle experience is inapplicable.

[go4mars]

... Perhaps a component of the ullage after separation can be flashed into compressed methane gas and oxygen gas within their respective systems by a light-weight, carefully controlled, exothermic reaction within the tanks to provide structural rigidity to the stage for re-entry (vapour pressure/fugacity), and force the ubiquitous presence of prop reagents through the plumbing as a gas for low initial thrust levels until the liquid component is forced back to the plumbing by acceleration of the stage (changing from gas propellants back to "the usual liquid burn")? 

 
It seems to me that this would work better with a staged combustion engine vs GG.  Is that right? 

No takers?  Well how about a couple related questions:

Has there ever been compressed gas rocket engines instead of liquid/solid rocket engines? non-orbital examples are fine.   

Has there ever been compressed methane and oxygen gas rocket engines?  Perhaps the expansion could even be a source of cooling (like how a fridge works)...

Would it damage any components to have some inefficient burps of compressed gases go through the plumbing of a falcon stage for a bit of thrust until the liquids settled back to where they should go? 


I just read that Buran used gaseous oxygen (GOX) for something but wikipedia didn't say what...   Anyone?

[Kaputnik]

Do you mean gas that has been compressed but not to the point of becoming liquid at any point on its way through the engine?
Wouldn't that greatly reduce the flow rates for a given size of piping/injector?

[baldusi]

LOX is 861 times more dense than the gaseous state at 1atm. Think of the implications of that.
Regardless, I think HMXHMX had a patent for a fully gaseous LAS system. In that particular case, the excess volume is not a problem, since it's still thinner than the capsule, and you save a lot of time in engine transients.

[Salo]

I just read that Buran used gaseous oxygen (GOX) for something but wikipedia didn't say what...   Anyone?

http://www.buran.ru/htm/odu.htm

[go4mars]

Do you mean gas that has been compressed but not to the point of becoming liquid at any point on its way through the engine?
Wouldn't that greatly reduce the flow rates for a given size of piping/injector?

  Yes and Yes.  This would entail relatively low rate gas and probably a lot of inefficiency for perhaps 10 seconds or so.  The turbo wouldn't come on until after the relevant parts of the system are liquid saturated again.  This is just to re-orient the small amount of liquid that is left for the boost-back burn within a mostly empty tank, and to provide internal pressure for structural strength on re-entry.  You would want the gaseous fluids to move out the back and up to the top of the tank so it didn't cause grief once the turbopump is flipped back on.   

[go4mars]

LOX is 861 times more dense than the gaseous state at 1atm. Think of the implications of that.

  The main implication is that if you only wanted the internal tank pressure at 1 atm (that would be 14.4 psi on the inside of the tank and a vacuum on the outside) then you would only have to flash a little bit of the liquid oxygen to gaseous state to increase the pressure a lot within the system.  IIRC, methane is more like 600 to 1.  So you would need to gasify a bit more of that, but still talking about very small fractions of the total.  The "thurst transient" on F1 flight 3 was residual of less than 10 psi with the "engine off".  But it packed a wallop.     

Regardless, I think HMXHMX had a patent for a fully gaseous LAS system. In that particular case, the excess volume is not a problem, since it's still thinner than the capsule, and you save a lot of time in engine transients.

  In this case, excess volume is not a problem either since there is a big empty tank to pressurize for re-entry, which doesn't require phase-changing very much of the LOX to achieve. 

[go4mars]

I just read that Buran used gaseous oxygen (GOX) for something but wikipedia didn't say what...   Anyone?

http://www.buran.ru/htm/odu.htm

Great source.  Thanks!  Google translate is wonderful. 

So it looks like this is a credible possibility for SpaceX after all, based on your link.  Some quotes (seems to be a few mis-translations in there denoted with a ? ): 

"money? supply propellants to the engine management, including the gasification of liquid oxygen "

"engine power control gasified with oxygen obtained in a special gas generator (gasifier) by combustion in oxygen a small fraction of the fuel"  -That would be the exothermic reaction I suggested...

A more interesting thing from their list of design decisions:  "intake of liquid fuel components under conditions close to weightlessness, with special fence devices based on fine-meshed (capillary) mesh blocks located in the lower parts of the tanks." 

If micro gravity capillary forces could keep only the liquid oxygen down by the tanks in the F9, then there wouldn't even necessarily need to be any gas down there.  The gas could be made (and kept) in the top half of the tank for pressure support and added structural rigidity for re-entry. 

The capillary force mesh would just need to be within the approximate area of tank levels at separation, and tank levels when they turn around to back into the parking lot (landing pad). 

Thanks again for the great link Salo! 

[baldusi]

Since I did my original numbers, it's been bothering me the fact that I did a same GLOW and T/W comparison, while SpaceX decisions appear to be heavily influenced by space constraints. Namely, the 3.65m width limitation. Thus, I decided to redo my numbers, but taking total volume in consideration. The method I used is crude in extreme, and I also did some parameter changes due to Shillings web's interface limitations. But as a rough idea it might lead somewhere. Shilling's forced changes is that since it only handles integer numbers, when I input 28.5 degrees from the Cape, it rounded it to 28, thus incurring in a "small" plane change maneuver. So I redid everything with 29 degrees. Same with the TLI of -1.8km^2/s^2 redefinition to -2.
To make a volume limited version, I first assumed a 465tonnes (with no payload) Falcon 9 v1.1. Then I assumed that a volume limited stage would have 1.3 less weight for a CH4 stage, and 3 times less weight for an H2 stage. I did kept the T/W fixed, thus, the engine thrust is different on most cases. This was a crude approximation, since I did no optimization on the First and Second stage's weight relationship. Also, I thought that the 88% mpf was too low for CH4, and thus increased it to 90%. I did calculate it with 88% and the results were not good. I'll be supplying them below.

Case	All RP-1	RP-1+H2	RP-1+CH4	All CH4	CH4+H2
T/W (No Payload)	1.225	1.225	1.225	1.225	1.225
GLOW (No Payload)	464,000	428,000	451,538	357,385	333,846
First Stage-Fuel	RP-1	RP-1	RP-1	CH4	CH4
Volume	1	1	1	1.3	1.3
First Stage-GLOW (kg)	408,000	408,000	408,000	313,846	313,846
First Stage-Fuel Fraction (%)	96.60%	96.60%	96.60%	95%	95%
First Stage-Dry Mass (kg)	13,872	13,872	13,872	15,692	15,692
First Stage-Fuel Mass (kg)	394,128	394,128	394,128	298,154	298,154
First Stage-Thrust (kN)	5,594	5,162	5,445	4,314	4,032
First Stage-ISP (vacuum sec)	310	310	310	355	355
Upper Stage-Fuel	RP-1	H2	CH4	CH4	H2
Volume	1	3	1.3	1.3	3
T/W (No Payload)	1.15	1.15	1.15	1.15	1.15
Upper Stage-GLOW (kg)	54,000	18,000	41,538	41,538	18,000
Upper Stage-Fuel Fraction (%)	92%	85%	90%	90%	85%
Upper Stage-Dry Mass (kg)	4,320	2,700	4,154	4,154	2,700
Upper Stage-Fuel Mass (kg)	49,680	15,300	37,385	37,385	15,300
Upper Stage-Thrust (kN)	631	225	491	491	225
Upper Stage-ISP (vacuum sec)	350	470	380	380	470
Fairing (kg)	2,000	2,000	2,000	2,000	2,000
Fairing Separation (s)	200	200	200	200	200
LEO: 200km x 29deg	15,578	15,175	16,001	15,421	15,159
GTO: 35786km x 185km x 29d	4,720	6,177	5,399	5,229	6,290
TLI: C3= -2 (185km x 29d)	2,940	4,672	3,631	3,505	4,778
TMI: C3= 13 (185km x 29d)	1,523	3,463	2,216	2,120	3,563
LEO with CH4 US pmf of 88%			15,060	14,513
GTO with CH4 US pmf of 88%			4,490	4,342
TLI with CH4 US pmf of 88%			2,752	2,629
TMI with CH4 US  pmf of 88%			1,343	1,254

When numbers are within a 10%, you can't really say much, but that they are within the margin of error. So, on principle, it would seem that's very hard to make a case for CH4 for LEO, when you are volume limited. It would give a significant improvement for higher energy orbits. But, it's almost the same than simply using a CH4 upper stage. The interesting point, is that even in this volume limited case, the H2 US is the superior choice! So, it would seem, when we take this new constraint into consideration, that an all CH4 doesn't necessarily makes so much extra sense. In fact, just doing a CH4 US gives the same or even better performance! Multiple factors have to be taken into consideration. So this is a first order approximation. What I would like to do next, is to assume a certain delta-v budget for the returning stage, plus adding some extra dry mass. And see how this plays out. What would be a reasonable delta-v budget for a first stage? 2.8km/s?

[Robotbeat]

Yeah, ch4 sort of sucks for a first stage if you are volume limited. If you subcool the kerosene to near methane temps (increasing its density further) you get significantly better performance for a first stage in a volume limited scenario.

[Ben the Space Brit]

@ baldusi,

Thanks for the work-out there; very informative.

I suspect that any LCH4 upper stage engine will be for commercial satellite and space probe launch only.  It would probably be quicker to market which is what SpaceX needs.  The LH2 high-energy upper stage (with that fantastic 150klbf/470s vacuum Isp) will probably need a few years work - too long to wait if SpaceX is going to challenge ILS and Arianespace for the comsat market.