AirLaunch Tests Propane-Fueled Rocket for Darpa Falcon Program
Aviation Week & Space Technology
09/05/2005, page 65
Michael A. Dornheim
Los Angeles
Self-pressurizing fuel a leading contender for Darpa Falcon
AirLaunch has finished initial tests of a propane-fueled uncooled rocket engine intended for the second stage of its planned QuickReach I launcher.
The work is for a Defense Advanced Research Projects Agency (Darpa) contract to reduce technical risk for its Falcon small launch vehicle (SLV) program. Four companies received these Phase 2A contracts in September 2004, which also required preliminary design of an SLV contender (AW&ST Sept. 27, 2004, p. 26). The agency plans to award one or two Phase 2B contracts by October to build and fly SLV prototypes by 2008.
AirLaunch chose propane to eliminate the expense of a high-pressure turbopump to fuel the engine, says Gary C. Hudson, the company founder. Instead, if propane is heated to 105F, it will self-pressurize the fuel tank to 200 psia., which is adequate to feed an upper-stage engine working against very low atmospheric pressure. The engine burns the fuel with liquid oxygen (LOX), which will self-pressurize to 200 psia. at about -250F, Hudson says. As propellants are consumed, their temperatures and pressures drop.
Methane is also self-pressurizing, but the propellant density of propane and LOX is about 10% more than methane/LOX, meaning smaller tanks for the propane system. Kerosene is denser but has no self-pressurizing capability.
The AirLaunch engine is designed to produce 24,000 lbf. thrust in vacuum with a high expansion ratio nozzle. The rocket shown here is a full-size combustion chamber and throat, with an outside diameter of about 18 in. The short nozzle is to prevent overexpansion for low-altitude tests at the Mojave, Calif., Civilian Flight Test Center. As of last week, the engine had under 20 sec. run time, but Hudson expects 50-100 more firings next year. Tests so far show the engine is stable and can be ignited quickly after shutdown, the company says. The tests are being conducted by HMX Inc., which Hudson co-founded in 1994.
The engine is made of composite materials, with an ablative liner to avoid the cost of liquid cooling, an approach taken by all Falcon SLV participants.
QuickReach I is to be dropped from a USAF/Boeing C-17 transport to assume a roughly vertical attitude and ignite for a midair launch (AW&ST June 27, p. 32).
I wonder if they could replace the LOX with nitrous oxide. You could eliminate the need for cryogenics altogether.
Still, LOX is cheap to come by.
It's good to see Gary Hudson back in the picture; Last I remember hearing about him, it was with the Roton and Phoenix. Zubrin said the Roton probably wouldn't work, but at least it was innovative.
Nitrous oxide has a lower boiling point than propane, so the vapour pressure for the same temperature should actually be higher. So from the pressure point of view at least, it should work.
But I suspect that the chemistry might not be as energetic, per unit mass, with N2O as with pure O2. Anybody care to guess what the main reaction products might be, besides the usual CO2 and H2O ?
It might work for a BDB (Big-Dumb-Booster). With few moving parts, you could churn out boosters much faster and for less than we do today. And given the scale of a BDB, you needn't worry about getting top-notch specific impulse: The payload will be huge in any case. But if AirLaunch can do it with LOX, there's no need for laughing gas in the first place. And as you have already stated, LOX is more energetic anyway.
Propane's chemical formula is: C3H8. Combusting it with nitrous oxide (N2O) yields:
C3H8 + 10N2O -----> 3CO2 + 4H20 +10N2
assuming perfect and complete combustion and of course no dissociation in rocket exhaust.
Nitrous oxide and Propane ought to burn just fine, but I don't know about expected Isp. The specific impulse (exhaust velocity in meters per second divided by the earth's gravitational acceleration of 9.80665 m/s^2) is a complex function of the molecular weight of exhaust products, the ratios of the enthalpy of the specific heats of the exhaust gasses at constant pressure to constant volume--which in of itself is complicated to determine as this depends upons the pressure too---and of course the enthalpy of combustion, the combustion chamber pressure (and this determines the operating temperature.) Basically, the only really good way to understand the performance of such a propellant combination is to build a small test motor and test it under many different conditions...
This can be done, but it is a major scientific undertaking all in itself.
Still, it could be done.
My guess is that the expected benefits of nitrous oxide would be overshadowed by the general availability of liquid oxygen, and the industry standard cryogenic valves, piping, etc. that is already available. Plus, liquid oxygen is a high density oxidizer possessing a density of nearly 1100 kilograms per cubic meter, which makes it even denser than water. This is a good trait to have in a rocket propellant.
Yes, well, I figured N2 might be fairly high on the list product species, but in a real rocket engine situation, are we talking well over 50%, or something less than that -- with the rest of the Ns tied up in stuff like NO, (CN)2, HCNO, etc. ?
Well...you're right, Jaro. You will end up with different products, which is why I worded the reply the way I did.
In actuality you end up with a whole array of different combustion product species, depending upon the actual combustion conditions. And there is a really nice program, developed by the US Air Force, which predicts the actual chemical composition of rocket exhaust, and is surprisingly available online 'free gratis' with no restrictions. I'll see if I can locate it--I'll post the links for it as soon as a re-find the program. It is simplicity itself to use.
The reality though is that the predominant exhaust products, greater than 95% will be composed of the primary exhaust products--i.e., those predicted by the Stoichiometric chemical equation I came up with in my previous post. The other 5% begin to have significant effects when trying to figure out the radiative heating of the inner chamber walls due to the presence of excited free radicals. In a large high performance rocket engine, radiative heating is very significant and estimates of this heat flux must be calculated before any metal is formed....
