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Hi to all.
In this patent
https://www.google.com/patents/US20140182265
It is written that a LOX-ammonia rocket with an expansion ratio of 100 can reach a specific impulse of 420 s, near to the 460 seconds of LOX-LH2 rockets but without the trouble of LH2 storage. Why this propellant combination was used so rarely?
Last edited by Quaoar (2016-12-14 16:33:26)
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http://www.astronautix.com/l/loxammonia.html
Vacuum Isp = 343s. About the same as LOX/kerosene. Density is about 80% of gasoline. Ammonia is toxic, which complicates things. On the plus side, ammonia is a much better coolant than a hydrocarbon, making it easier to design high pressure engines. Nitrogen is not a greenhouse gas, although water vapour is. On balance, ammonia would however appear to be more climate friendly than kerosene. On the other hand, it is highly toxic to marine life.
Last edited by Antius (2016-12-14 18:39:26)
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The burning would yield 4 NH3 + 5 O2 = 6 H2O + 4 NO as a reaction with the NO being also toxic...
Now for a lunar launch fuel there is not any issue as there is not any chance to be exposed to it and while it is the same for mars it would however be an atmospheric builder ....
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http://www.astronautix.com/l/loxammonia.html
Vacuum Isp = 343s. About the same as LOX/kerosene. Density is about 80% of gasoline. Ammonia is toxic, which complicates things. On the plus side, ammonia is a much better coolant than a hydrocarbon, making it easier to design high pressure engines. Nitrogen is not a greenhouse gas, although water vapour is. On balance, ammonia would however appear to be more climate friendly than kerosene. On the other hand, it is highly toxic to marine life.
That was the XLR-99 of the X-15 spaceplane, which was projected for working in high atmosphere and don't have a high expansion ratio. I guess that it would have a quite better Isp with a bigger nozzle.
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Would that work with NTO in place of LOX, thereby getting rid of the need for cryogenic tankage?
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Would that work with NTO in place of LOX, thereby getting rid of the need for cryogenic tankage?
It could be interesting, but I didn't find any data on NTO-NH3 rockets.
Anyway, my interest in ammonia rocket is due to the spaceship I would like to convince for my next hard-sf novel: she has an Orion-drive NNP, so she needs to store a lot of ammonia for running the shock absorbers and the cannon for the pulse units. Using ammonia even as a propellant for the RCS would simplify the plumbings. LOX is also necessary as a reservoir of breathing gas for passengers and crew. So that's why I thought about LOX-NH2 rockets.
Last edited by Quaoar (2016-12-15 11:09:53)
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I don't have any data available, but what about B2H6? Commonly called Diborane. This at one time, was thought to have some promise.
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The XLR-99’s that pushed the X-15’s used LOX-NH3 propellants. I do not know the expansion ratio of those engines. What works depends upon the ratio of chamber pressure to ambient pressure at the lowest altitude the engine is designed for.
My ancient Pratt and Whitney Aeronautical Vest Pocket Handbook from Dec 1969 lists data for this propellant combination, among many others. It seems to be as accurate a reference today as it was back then. Isp depends marketedly upon the nozzle expansion conditions, c* depends only upon chamber pressure, and so is a less confusing measure for comparison; although, you have to do your own ballistics (at least in rudimentary form) to get Isp from it via Isp = CF c* /gc. It is CF that is dependent upon the expansion pressure, geometric area ratio, and exit divergence angle (which correction applies only to the momentum terms, not the pressure-area terms).
Quoting from my reference at two chamber pressures with different expansion conditions:
Pc, psia r Tc, R c*, fps Isp, s how expanded
1000 1.41 5580 5880 295 note 1
100 1.41 5260 5790 347 note 2
Note 1: perfect expansion to sea level 14.7 psia, shifting equilibrium
Note 2: perfect expansion into vacuum at Ae/At = 40, shifting equilibrium
“perfect expansion” means using a nozzle kinetic energy efficiency = 1, when it is usually near 0.983
These values more-or-less bracket the practical achievable values; although Ae/At > 40 is possible
Myself, I would use a power-function curve-fit for c* = (5880 ft/sec) (Pc, psia / 1000)^0.006699 from these data. And I would do my own CF calculation more-or-less the way the textbooks have it for my own expansion conditions and calculations, but being very careful to apply my nozzle efficiency only to the momentum terms. Not all the textbooks do this efficiency thing quite correctly, so beware.
I would use a gas specific heat ratio near 1.2 for this, in the absence of any better data. I would use efficiency = 0.5*(1 + cosine(half-angle)) for my nozzle kinetic energy efficiency. Half angle is the conical half angle if a conical nozzle, and is the average of the near-throat and exit half-angles if a curved expansion bell. That ought to get you within a tiny fraction of a percent the right value to use.
That’s just standard real-world ballistics stuff. I used to do this crap (and much more besides) for a living. Solids, liquids, hybrids, ramjet nozzles-only, etc, all the same, no difference. It starts with c* at your chamber pressure.
GW
Last edited by GW Johnson (2016-12-15 16:03:22)
GW Johnson
McGregor, Texas
"There is nothing as expensive as a dead crew, especially one dead from a bad management decision"
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Here are some of the other topics that we talk about Diborane in.
LH2 fuel replacement... - ...NaH / NaBH4?
Silane Hoppers - Use the CO2 man...
LOX/CO2-Diborane rocket for reusable landing boats
The trick is too use what we can get for the least amount of energy to make it....
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The XLR-99’s that pushed the X-15’s used LOX-NH3 propellants. I do not know the expansion ratio of those engines. What works depends upon the ratio of chamber pressure to ambient pressure at the lowest altitude the engine is designed for.
My ancient Pratt and Whitney Aeronautical Vest Pocket Handbook from Dec 1969 lists data for this propellant combination, among many others. It seems to be as accurate a reference today as it was back then. Isp depends marketedly upon the nozzle expansion conditions, c* depends only upon chamber pressure, and so is a less confusing measure for comparison; although, you have to do your own ballistics (at least in rudimentary form) to get Isp from it via Isp = CF c* /gc. It is CF that is dependent upon the expansion pressure, geometric area ratio, and exit divergence angle (which correction applies only to the momentum terms, not the pressure-area terms).
Quoting from my reference at two chamber pressures with different expansion conditions:
Pc, psia r Tc, R c*, fps Isp, s how expanded
1000 1.41 5580 5880 295 note 1
100 1.41 5260 5790 347 note 2
Note 1: perfect expansion to sea level 14.7 psia, shifting equilibrium
Note 2: perfect expansion into vacuum at Ae/At = 40, shifting equilibrium
“perfect expansion” means using a nozzle kinetic energy efficiency = 1, when it is usually near 0.983
These values more-or-less bracket the practical achievable values; although Ae/At > 40 is possibleMyself, I would use a power-function curve-fit for c* = (5880 ft/sec) (Pc, psia / 1000)^0.006699 from these data. And I would do my own CF calculation more-or-less the way the textbooks have it for my own expansion conditions and calculations, but being very careful to apply my nozzle efficiency only to the momentum terms. Not all the textbooks do this efficiency thing quite correctly, so beware.
I would use a gas specific heat ratio near 1.2 for this, in the absence of any better data. I would use efficiency = 0.5*(1 + cosine(half-angle)) for my nozzle kinetic energy efficiency. Half angle is the conical half angle if a conical nozzle, and is the average of the near-throat and exit half-angles if a curved expansion bell. That ought to get you within a tiny fraction of a percent the right value to use.
That’s just standard real-world ballistics stuff. I used to do this crap (and much more besides) for a living. Solids, liquids, hybrids, ramjet nozzles-only, etc, all the same, no difference. It starts with c* at your chamber pressure.
GW
Thanks GW,
347 s is quite good. Do you think that LOX-NH3 rockets may be a good choice for the RCS of an Orion-drive spaceship, which needs ammonia anyway for the pulse-unit cannon and for the shock-absorbers?
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NH3-LOX:
Well, performance is good, but it's not hypergolic. You have to light off the combustion with a torch or pyrotechnic every time you fire the thing.
Diboranes, etc:
Again looking in my old Pratt & Whitney handbook, I only see B2H6 diborane (and also B5H9 pentaborane) listed as a fuel with F2 oxidizer, plus listings with OF2 and H2O2 for diborane only. In the 1960’s fluorine F2 was tried experimentally, but found to be way too toxic and too corrosive to contemplate further.
Same definitions as my previous posting: 1000 psia Pc into atmospheric, no losses; and 100 psia into vacuum at Ae/At = 40, no losses. Isp in seconds:
F2 B2H6 371 SL 431 vac
F2 B5H9 360 SL 417 vac
OF2 B2H6 372 SL 434 vac
H2O2 B2H6 333 SL 391 vac
What that means is the only combination not involving fluorine is hydrogen peroxide, which as a rocket-grade product is 90 to 95% H2O2, with the rest water. That stuff is technically storable, but it is very unstable. It spontaneously explodes in extremely-violent decomposition, after an erratic and unpredictable interval measured in days. Minutes, if you have a flaw or contaminant particle in your storage tank. You use it distilled-up to strength only hours before use. To store safely, it must be under 50%, as a generally-recognized industrial safety practice.
If you look instead at LOX and methane, or RP-1 kerosene, the same reference lists:
O2 CH4 310 SL 365 vac
O2 RP-1 299 SL 351 vac
I really don’t see much Isp advantage to justify the handling and toxicity risks of the diborane itself, myself. H2O2-B2H6 is only slightly better than O2-CH4, and only somewhat better than O2-RP-1. And you still have the H2O2 explosion risk, while LOX-methane is just a couple of mild cryogens.
There’s even LOX-hydrazine, although I don’t remember this ever actually being flown:
O2 N2H4 313 SL 367 vac
And the reference “best” is of course:
O2 H2 388 SL 454 vac
Theoretically, you might get away with trying to handle ozone, I suppose. Although no one has ever flown such a thing, and probably for good reasons. It has the highest Isp numbers of anything, including all the fluorine options. Its toxicity is listed as “high” just like all the fluorine compounds. In comparison, the toxicity of ammonia (which danger we all understand) is listed as “medium”. Which means in practical terms that ozone is truly horrendous, just like the fluorines. Same is true for the “high” toxicity boranes.
O3 H2 423 SL 493 vac
All in all, I think we’re far better off just trying to handle LOX, kerosene, liquid methane, hydrogen, and (for lower Isp thrusters) NTO and MMH or UDMH. Sorry. NTO-hydrazine (N2H4) outperforms IRFNA and kerosene, the other storable. IRFNA is toxic (“medium”) and corrosive (one shot only, and fuel it up right before you fly). All three are hypergolic ignition.
NTO-hydrazine 292 SL 342 vac
NTO-UDMH 287 SL 336 vac
IRFNA-RP-1 263 SL 309 vac
GW
Last edited by GW Johnson (2016-12-16 13:36:05)
GW Johnson
McGregor, Texas
"There is nothing as expensive as a dead crew, especially one dead from a bad management decision"
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Stabilised H2O2 gets shipped around in tank cars at 70% for the pulp industry. It was shipped from Sydney to Woomera at 86% (I understand) for rocket propellant (with kerosene). Stabilisers must be good enough. Even 70% peroxide would convert itself entirely to Oxygen and steam.
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I don't know anything about stabilizers for hydrogen peroxide, but then I am an old obsolete retiree. The last info I had about stateside usage is ~7 years old, and that's storage at 50%, distill up for use to 90-95%, only hours before use.
You monitor fluid temperatures in the peroxide tank over time. If you see a rise, dilute below 50% immediately. The last occupant of the old Texas rocket plant where Spacex is now, was Beal Aerospace, who did H2O2-kerosene rocket work. They violated the monitoring rule, and that tank of peroxide blew up, causing a lot of expensive damage.
GW
GW Johnson
McGregor, Texas
"There is nothing as expensive as a dead crew, especially one dead from a bad management decision"
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