Quote from: kevin-rf on 10/10/2014 05:42 pmIf the other two engine families where not also fist stage engines, I would say it had something to do with a sealevel optimized engine, but that is not the case here. Look beyond the engines to the LVs which use them. Ariane and H-II have solid boosters which do most of the lifting at low altitudes. Thus you don't have to worry about core engine sea level Isp so much. Situation is different with Delta IV, good sea level performance is required from RS-68.sea level Isps:Vulcain 1: 326sVulcain 2: 318sLE-7: 349sLE-7A: 338sRS-68: 365s
If the other two engine families where not also fist stage engines, I would say it had something to do with a sealevel optimized engine, but that is not the case here.
Just a guess? Thrust. The low area ratio implies a big throat. Have you compared the propellant mass flow of the engines?
Remember the RS-68 was designed for cost. Hence the use of an ablative nozzle instead of a regeneratively cooled nozzle. I wonder if it had something to do with manufacturing costs. If the other two engine families where not also fist stage engines, I would say it had something to do with a sealevel optimized engine, but that is not the case here. Could it be the weight of a larger nozzle would have negated the ISP gain?
>>mocular mass ??It should molecular mass Or, better yet - molar mass.The answer to your question is in chemical thermodynamics. Which is bad, because in general opinion this is very complex matter. But the good news is that here you need only the simplest part of all the thermodynamics - its First Law and the term Enthalpy.The answer - in a short - LH2 has very high combustion Enthalpy (which is here energy per mole) AND the lowest possible molar mass.This gives to the LH2-LOX pair the highest specific combustion energy -- per gram of propellant - among all the things which can burn (in the Universe:)
Larger exit area means larger value for the "(Pe-Pa)Ae/q" part of Isp.
As for the thermodynamics, one doesn't have to understand thermodynamics thoroughly to design rocket engines, what he need to do is learn to read relative charts.
Quote from: nimbostratus on 10/11/2014 09:12 amLarger exit area means larger value for the "(Pe-Pa)Ae/q" part of Isp.Too simplified statement. For instance engine optimized for vacuum should have as low Pe as possible. Play with engine simulation tools, the greater the area ratio and the lower the Pe the closer optimum expansion and vacuum Isps are, meaning pressure thrust component shrinks.Optimum would be not to have the pressure thrust component at all but in practice you always have some. At sea level it is usually negative, meaning optimum expansion doesn't happen at sea level but at higher altitude (~4-5km for many kerolox, even higher for hydrolox core engines you listed (and SSME)).
I can't say you are wrong, you just missed the point.I was not talking about optimizing a liquid hydrogen engine, but comparing liquid hydrogen engines with other chemical rocket engines.
For the "(Pe-Pa)Ae/q" part, low average molecular mass means larger exit area even for the same thrust and optimzed for the same ambient pressure, which can be seen as in the case of rd-0120 and rd-170.
At lease one thing is for sure, the optimized ambient pressure for rd-0120 and rd170 is approximately sea level pressure.
Quote from: nimbostratus on 10/11/2014 02:57 pmAt lease one thing is for sure, the optimized ambient pressure for rd-0120 and rd170 is approximately sea level pressure.Incorrect. ~0.5atm for RD-170 and ~0.16atm for RD-0120.
Quote from: R7 on 10/11/2014 03:01 pmIncorrect. ~0.5atm for RD-170 and ~0.16atm for RD-0120.Can't believe that.
Incorrect. ~0.5atm for RD-170 and ~0.16atm for RD-0120.
Quote from: nimbostratus on 10/11/2014 03:20 pmQuote from: R7 on 10/11/2014 03:01 pmIncorrect. ~0.5atm for RD-170 and ~0.16atm for RD-0120.Can't believe that.It's not a matter of belief, that's how it is.
Can't believe that.
Quote from: nimbostratus on 10/11/2014 03:20 pmCan't believe that.An understanding of thermodynamics would open your eyes.
OK,and thanks for sharing these sceenshots.It seems that you have very good simulation software.Can you also provide the exit area of a chamber of rd-170 and that of rd-0120 optimized for a common ambient pressure, say 0.5atm?
Quote from: nimbostratus on 10/11/2014 04:24 pmOK,and thanks for sharing these sceenshots.It seems that you have very good simulation software.Can you also provide the exit area of a chamber of rd-170 and that of rd-0120 optimized for a common ambient pressure, say 0.5atm?You are welcome. It is a good software indeed, developed by a real Russian rocket scientist Alexander Ponomarenko. Go ahead and try it yourself. RD-170 configuration file is included in the software examples, RD-0120 you can quickly build from public information but I'll attach configuration file for that to speed things up. Remove the ".txt" from the filename, forum software requires it.
@nimbostratusYou are not supposed to read the attached file but load it into the software I provided link for Chamber temperature is connected to chamber pressure. RD-170 Pc is very high, therefore the Tc is high too. IIRC only RD-180 has higher Pc among kerolox engines, the RPA tool calculates Tc nearly 3900K for it.