Quote from: Proponent on 02/07/2013 09:21 amI'm still wondering why methane seems to be the clear favorite, when it's so much less dense than other hydrocarbons. Even if ISRU methane is used as a fuel on Mars someday, that's some distance into the future and in the meantime an awful lot more stuff is going to be and will continue to be launched from Earth than from Mars. Justifying methane over the others on this basis seems to be a case of the tail wagging the dog.Methane, with its simple C-H bonds, probably is less subject to coking, but is coking really a significant problem with the others? Surely there must be some information out there about coking, like reaction coefficients for polymerization as a function of temperature for the various fuels. This would reduce the arm-waviness of the discussion.It's not just the coking, while that is nice. Methane rich gas has a very high specific heat. This means that for a fixed turbine inlet temperature, you can get ungodly amounts of power out of it.This is why methane, like hydrogen, optimizes at a fuel-rich preburner for staged combustion. The lack of coking just helps close the case.Allows you either to have a very low turbine temp and get a moderate chamber pressure, which is good for reusability, or a very high chamber pressure for a typical turbine inlet temp (900-1200K), which is good for performance.
I'm still wondering why methane seems to be the clear favorite, when it's so much less dense than other hydrocarbons. Even if ISRU methane is used as a fuel on Mars someday, that's some distance into the future and in the meantime an awful lot more stuff is going to be and will continue to be launched from Earth than from Mars. Justifying methane over the others on this basis seems to be a case of the tail wagging the dog.Methane, with its simple C-H bonds, probably is less subject to coking, but is coking really a significant problem with the others? Surely there must be some information out there about coking, like reaction coefficients for polymerization as a function of temperature for the various fuels. This would reduce the arm-waviness of the discussion.
I'm not sure I understand the rationale for using the ratio of propellant volume to burn-out mass as a metric. If I'm using a bulky propellant combination like lox-hydrogen, I'm going to tend to have large, heavy tanks, which is bad. But since the tank mass appears in the denominator, I'm some sense the combination is rewarded for being bulky.
Here's a graph showing how poorly methalox performs. We plot delta-V versus the ratio of propellant volume to final mass. Two lowest curves are hydrolox with a nominal mixture ratio (MR) of 6 and an impractical one of 7.5 (8 is stoichiometric). The next worst is methalox. All the other combinations perform better.Say for example you want your first stage to have a 4 km/s delta-V, about what you need to get to LEO for the first stage of a two vehicle with the same propellants. Hydrolox requires 4 litres of propellant for every kg of your total burnout mass (which includes the first stage dry mass, second stage and payload). Methalox requires 2.35 L/kg. Kerolox requires 2.0 L/kg. That is, your first stage needs 18% more propellant volume which corresponds to about 18% more propellant tank mass.
Quote from: Steven Pietrobon on 10/24/2014 08:20 amHere's a graph showing how poorly methalox performs. We plot delta-V versus the ratio of propellant volume to final mass. Two lowest curves are hydrolox with a nominal mixture ratio (MR) of 6 and an impractical one of 7.5 (8 is stoichiometric). The next worst is methalox. All the other combinations perform better.Say for example you want your first stage to have a 4 km/s delta-V, about what you need to get to LEO for the first stage of a two vehicle with the same propellants. Hydrolox requires 4 litres of propellant for every kg of your total burnout mass (which includes the first stage dry mass, second stage and payload). Methalox requires 2.35 L/kg. Kerolox requires 2.0 L/kg. That is, your first stage needs 18% more propellant volume which corresponds to about 18% more propellant tank mass.Musk has confirmed that his methalox will be sub-cooled close to freezing temps. How does that affect the density and other properties (EG having to add more heat to reach the same combustion temps [impact to Isp?], reduced energy to pump a smaller volume, viscosity effects, extra energy required to autogenously pressurise)? Cheers, Martin
In Steven's formulation, the best propellants that have ever flown extensively are hypergols...However, I'm not sure that volume is the correct normalization here, as tank mass much more closely scales to surface area, which scales as volume^(2/3). If you apply that correction, the more exotic dense fuels will appear not much better than methane.
Quote from: MP99 on 10/27/2014 08:18 amQuote from: Steven Pietrobon on 10/24/2014 08:20 amHere's a graph showing how poorly methalox performs. We plot delta-V versus the ratio of propellant volume to final mass. Two lowest curves are hydrolox with a nominal mixture ratio (MR) of 6 and an impractical one of 7.5 (8 is stoichiometric). The next worst is methalox. All the other combinations perform better.Say for example you want your first stage to have a 4 km/s delta-V, about what you need to get to LEO for the first stage of a two vehicle with the same propellants. Hydrolox requires 4 litres of propellant for every kg of your total burnout mass (which includes the first stage dry mass, second stage and payload). Methalox requires 2.35 L/kg. Kerolox requires 2.0 L/kg. That is, your first stage needs 18% more propellant volume which corresponds to about 18% more propellant tank mass.Musk has confirmed that his methalox will be sub-cooled close to freezing temps. How does that affect the density and other properties (EG having to add more heat to reach the same combustion temps [impact to Isp?], reduced energy to pump a smaller volume, viscosity effects, extra energy required to autogenously pressurise)? Cheers, MartinJust subcooling the CH4 (which you get "for free" with a small common bulkhead), gives less than 3% improvement in propellant mass for same volume. Doing full CH4@93K and LOX@68K is a little better 8.5%. It might not seem that much, but this is the rough performance improvement expected:Densification\OrbitLEOGTOLOX+CH4+15.00%+23.00%CH4 Only+6.50%+10.50%Which is quite interesting if you ask me. It is roughly like adding two solids to an EELV, for example. And almost like the RS-68 to RS-68A improvement on the Delta IV Heavy. Say that your rocket does 5.3 tonnes to GTO, full densification would bring it to 6.5 tonnes. And CH4 would allow 5.85 tonnes. So, for cases where you are a bit below your target performance, you could apply this and get an extra decade out of your design. Or save this as an option in design and have some margin for any other performance shortcoming that you might have.
Quote from: strangequark on 02/07/2013 03:42 pmQuote from: Proponent on 02/07/2013 09:21 amI'm still wondering why methane seems to be the clear favorite, when it's so much less dense than other hydrocarbons. Even if ISRU methane is used as a fuel on Mars someday, that's some distance into the future and in the meantime an awful lot more stuff is going to be and will continue to be launched from Earth than from Mars. Justifying methane over the others on this basis seems to be a case of the tail wagging the dog.Methane, with its simple C-H bonds, probably is less subject to coking, but is coking really a significant problem with the others? Surely there must be some information out there about coking, like reaction coefficients for polymerization as a function of temperature for the various fuels. This would reduce the arm-waviness of the discussion.It's not just the coking, while that is nice. Methane rich gas has a very high specific heat. This means that for a fixed turbine inlet temperature, you can get ungodly amounts of power out of it.This is why methane, like hydrogen, optimizes at a fuel-rich preburner for staged combustion. The lack of coking just helps close the case.Allows you either to have a very low turbine temp and get a moderate chamber pressure, which is good for reusability, or a very high chamber pressure for a typical turbine inlet temp (900-1200K), which is good for performance.This has seemed to me to be the strongest argument for methane. But I've been thinking about it a little more.My earlier post containing heat capacities of light hydrocarbons shows that methane's is a bit higher than those of other light hydrocarbons. Thus, at a given temperature, methane packs somewhat more thermal energy for running a turbopump. That's obviously good.But... that energy is used to pump propellants, and the power required by a pump depends on the volume rate that's pumped, not on the mass rate. So, let's compute the heat capacity per unit volume of propellant (see the third attachment for the calculations). The results are plotted below, with underlying data from the NIST Chemisty WebBook. The first plot shows heat capacity of the fuel per unit volume of propellant for hydrogen at O/F=5.5, methane at 3.5, ethane at 3.2, ethylene (ethene) at 2.6, propane at 3.9, and propylene (propene) at 2.7. This figure is meant to represent fuel-rich staged combustion. The second plot is the same except that the heat capacity of oxygen is added in, corresponding to full-flow staged combustion, where the temperatures at the inlets of the two turbines are the same.To make visual sense of the plots, note that deeply-cryogenic hydrogen is plotted in the coldest color, blue. The colors for the hydrocarbons will make sense if you know the resistor color code; brown = 1 (carbon atom), red = 2, orange = 3.In FRSC at 700 K, the hydrocarbon to beat is propane, with a heat capacity per unit volume of propellant of 740 kJ K-1 m-3. Methane comes in about 10% lower at 670 kJ K-1 m-3.Propane also comes out tops In FFSC at 700 K, with a heat capacity per unit volume of propellant of 1450 kJ K-1 m-3. Methane at 1340 kJ K-1 m-3 is several percent lower and is the worst of hydrocarbons considered here.Fold in methane's disadvantage in bulk density (830 kg/m3 vs. 920 kg/m3 for propane), and its few seconds' worth of Isp advantage over propane (and disadvantage in comparison to propylene) doesn't seem worth it, especially for a booster stage.Since the dudes at SpaceX (FFSC) and Blue Origin (FRSC) are smart and know a lot more about rocket engines than I do, I'm sure there are good reasons for preferring methane over other light hydrocarbons, but it doesn't look to me like heat capacity is one of them.
I assumed the output of the fuel-rich preburner would be a lot like the fuel, since only a small amount of the fuel is burned in the preburner (that has to be the case, otherwise staged combustion would not be efficient).
Quote from: Proponent on 10/27/2014 10:30 pmI assumed the output of the fuel-rich preburner would be a lot like the fuel, since only a small amount of the fuel is burned in the preburner (that has to be the case, otherwise staged combustion would not be efficient).It is but incomplete combustion means the rest is quite messy soup. Combustion of methane happens in numerous steps and it is a lottery how far each methane molecule gets in the intermediate steps before there's no more oxygen to drive the reaction to final products, water and CO2.
A question follows. Could a vehicle with these higher impulse fuels be launched from Earth benefiting from the higher impulse efficiency and be in situ refueled on Mars using old fashioned methane? The same engine with two different fuels.
Steven Pietrobon: I think Zubrin himself has fully acknowledged that Ethylene is completely superior to Methane and if he had the whole thing to do over again he would have pushed that instead as it's synthesis is almost as easy as methane, higher hydrocarbons not so much.Lower hydrogen needs for Ethylene and easier refrigeration (practically none on Mars) are considered even more important then the density and impulse values. The only reason to go for Methane now is that fact that everyone is developing LNG based engines for launch vehicles now and you could reuse thouse engines on Mars, but even then I suspect a dual fuel engine would be possible and advantageous.
