Quote from: Robotbeat on 02/06/2013 05:35 pmQuote from: R7 on 02/06/2013 05:34 pmQuote from: Robotbeat on 02/06/2013 05:28 pm(And of course the first one is talking about SSTO with the volume of the craft remaining constant.)Which is the point. Switching to methane does not help, unless you design a bigger vehicle to go with it.No, the conclusions only apply to SSTO and somewhat to first stages (according to your second one).Again, your second link said this:"So Impulse Density doesn’t mean a whole lot for upper stages that don’t interact with the atmosphere, but it does have a little bit of meaning when applied to boosters. For a booster stage, higher impulse density can result in a heavier, but smaller vehicle (generally)."R7, attached is a graph that shows the relative increase in Isp required to counteract the performance decrease due to density decrease. Rho_t is the mass of stage required to enclose a unit volume of propellant. Centaur is about 25, Atlas is around 70, Delta is around 50, Zenit is around 100, as a framework (kg/m^3). I would be happy to discuss assumptions, derivation, and sensitivity later tonight, but we should split the thread if so.
Quote from: R7 on 02/06/2013 05:34 pmQuote from: Robotbeat on 02/06/2013 05:28 pm(And of course the first one is talking about SSTO with the volume of the craft remaining constant.)Which is the point. Switching to methane does not help, unless you design a bigger vehicle to go with it.No, the conclusions only apply to SSTO and somewhat to first stages (according to your second one).Again, your second link said this:"So Impulse Density doesn’t mean a whole lot for upper stages that don’t interact with the atmosphere, but it does have a little bit of meaning when applied to boosters. For a booster stage, higher impulse density can result in a heavier, but smaller vehicle (generally)."
Quote from: Robotbeat on 02/06/2013 05:28 pm(And of course the first one is talking about SSTO with the volume of the craft remaining constant.)Which is the point. Switching to methane does not help, unless you design a bigger vehicle to go with it.
(And of course the first one is talking about SSTO with the volume of the craft remaining constant.)
I'll comment about the methane coking here too. AFAIK it's practically non-coking. Guessing this is related to problems in petrochemistry in converting methane to more useful things like ethylene. Those C-H bonds are really strong, search for efficient conversion catalysts/processes is on.
Quote from: R7 on 02/06/2013 08:43 pmI'll comment about the methane coking here too. AFAIK it's practically non-coking. Guessing this is related to problems in petrochemistry in converting methane to more useful things like ethylene. Those C-H bonds are really strong, search for efficient conversion catalysts/processes is on.Sorry to already take you off topic but does that mean that ISRU produced methane may not be as good a feedstock for further chemical reactions (looking WAY out at a Mars colony that wants to make some plastics) as, say, ethanol? (another possible feedstock that a colony may have access to if it has lots of algae tubes or similar)
Of course, you'd probably /start out/ with hydrogen and carbon monoxide when doing ISRU on Mars, so it's probably best to skip the middle man.
What are the assumptions and equations behind the graph?
Methane is *extremely* cheap and abundant.
[From the previous thread.]Quote from: Funchucks on 02/06/2013 06:13 pmMethane is *extremely* cheap and abundant.LNG is cheap and abundant, but pure methane is another matter. And the costs of LNG, LPG (propane plus impurities, largely butane) and propylene are all within about a factor of two of each other. Don't know about methyl acetylene. Since propellant costs are just a fraction of a percent of total launch costs, any price differences among the hydrocarbon fuels not only don't matter, but won't matter for a long time.I would guess that of these three, propylene is the easiest to obtain in relatively pure form (it's used in large quantities as a feedstock for producing plastics).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.
The are no methane/LOX engines that had been flown (AFAIK)....
... so there is still a lot of things to do and test before they become operational. So in other words it will still take some time.ISRU and methane production on Mars will also take some time, as you have mentioned. So at current moment in time you (as a engine designer/manufacturer) are faced with following problem: are you going to develop one engine type for launching from Earth (non-methane) and another one for launching from Mars (methane) or do you develop a single type of technology for both (methane)?At the moment, if you follow second option, you are limited to methane, as it is the most viable solution for Mars, that is also usable on Earth. If you could find other type of fuel that could be easily obtained/manufactured on both Earth and Mars and it would be usable in both locations, then you could substitute methane with it.
R7, you used an awesome Isp for kerosene, and a terrible one for methane. You're comparing apples and oranges. The RD-0124 has the best Isp of any kerosene engine on the planet at 359 sec. A methane-lox engine, as modeled using the Lewis code, under a frozen flow assumption (which is conservative and underpredicts), using the same chamber pressure and same expansion ratio (82) will get 382 sec.Difference is 6.4% For a light upper-stage (rho_t of 25, comparable to Centaur), you only need 5.3% to break even.The advantage is actually more pronounced for a first stage, a methane equivalent to the RD-180 beats it by 10.5%, when it only needs 7.8% to break even at rho_t of 100.
Btw found great paper on effects of impulse density http://www.dtic.mil/dtic/tr/fulltext/u2/283940.pdf, gotta read this with time.
Quote from: luksol on 02/07/2013 09:45 amThe are no methane/LOX engines that had been flown (AFAIK)....That's substantially true, but just to indulge my love of trivia, let me point out two flown lox-methane engines I'm aware of. These are the first liquid-fuel rocket in Europe and those by the CALVEIN group. CALVEIN has also flown lox-propylene.
Quote... so there is still a lot of things to do and test before they become operational. So in other words it will still take some time.ISRU and methane production on Mars will also take some time, as you have mentioned. So at current moment in time you (as a engine designer/manufacturer) are faced with following problem: are you going to develop one engine type for launching from Earth (non-methane) and another one for launching from Mars (methane) or do you develop a single type of technology for both (methane)?At the moment, if you follow second option, you are limited to methane, as it is the most viable solution for Mars, that is also usable on Earth. If you could find other type of fuel that could be easily obtained/manufactured on both Earth and Mars and it would be usable in both locations, then you could substitute methane with it.It still seems to me that the argument for methane now basically boils down to believing that large-scale methane production on Mars is imminent.
Even if SpaceX were to reach Mars in the time frame sometimes mentioned by Musk, it remains the case that:1. Between now and then, and for a long time afterwards, the tonnage launched from Earth will exceed the tonnage launched from Mars by orders of magnitude, and2. Setting up ISRU on Mars will require a huge amount of R&D anyway; adapting a non-methane-light-hydrocarbon engine for methane will be trivial in comparison, especially if the possible need for methane in the future is borne in mind (after all, the RL-10 has been run on methane without much modification, and that's a much larger change of propellant).Hence, I still don't see how possible future martian ISRU has carries much weight in selecting propellants now. It may even turn out that carbon monoxide will be the ISRU fuel of choice on Mars (the Isp isn't great, but in martian gravity it doesn't matter so much, and it's a lot easier to make). For that matter, what about lox-hydrogen? Getting the hydrogen is the hard part about martian methane production. If it turns out that the easiest way to get hydrogen on Mars to crack polar water, then maybe lox-hydrogen starts looking better than lox-methane for ISRU. Super Isp and drag and gravity losses associated with a low-density, high-Isp fuel won't be so significant in Mars's thin atmosphere and low gravity.If the choice between methane and another light hydrocarbon is just about a wash, as it may be for some systems, then I can see letting possibilities for martian ISRU be the deciding factor.
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.
Future high chamber pressure LOX/hydrocarbon booster engines require copper-base alloy main combustion chamber coolant channels similar to the SSME to provide adequate cooling and resuable engine life. Therefore, it is of vital importance to evaluate the heat transfer characteristics and coking thresholds for LNG (94% methane) cooling, with a copper-base alloy material adjacent to the fuel coolant. High-pressure methane cooling and coking characteristics were recently evaluated using stainless-steel heated tubes at methane bulk temperatures and coolant wall temperatures typical of advanced engine operation except at lower heat fluxes as limited by the tube material. As expected, there was no coking observed. However, coking evaluations need be conducted with a copper-base surface exposed to the methane coolant at higher heat fluxes approaching those of future high chamber pressure engines.
Hence, I still don't see how possible future martian ISRU has carries much weight in selecting propellants now. It may even turn out that carbon monoxide will be the ISRU fuel of choice on Mars (the Isp isn't great, but in martian gravity it doesn't matter so much, and it's a lot easier to make). For that matter, what about lox-hydrogen? Getting the hydrogen is the hard part about martian methane production. If it turns out that the easiest way to get hydrogen on Mars to crack polar water, then maybe lox-hydrogen starts looking better than lox-methane for ISRU. Super Isp and drag and gravity losses associated with a low-density, high-Isp fuel won't be so significant in Mars's thin atmosphere and low gravity.
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.
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.
Speaking of methane fueled engines, what happened to C&Space? IIRC South Korean company, faint memory that they testfired rmethane engine, good enough for smaller LV booster.
Quote from: R7 on 02/07/2013 12:11 pm, what happened to C&Space? IIRC South Korean companyA vague memory says that outfit moved to the US.
, what happened to C&Space? IIRC South Korean company
LNG is cheap and abundant, but pure methane is another matter.
1) And the costs of LNG, LPG (propane plus impurities, largely butane) and propylene are all within about a factor of two of each other. 2) Since propellant costs are just a fraction of a percent of total launch costs, any price differences among the hydrocarbon fuels not only don't matter, but won't matter for a long time.
Even if ISRU methane is used as a fuel on Mars some day, 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...
the tonnage launched from Earth will exceed the tonnage launched from Mars by orders of magnitude,
2. Setting up ISRU on Mars will require a huge amount of R&D
It may even turn out that carbon monoxide will be the ISRU fuel of choice on Mars (the Isp isn't great, but in martian gravity it doesn't matter so much, and it's a lot easier to make). For that matter, what about lox-hydrogen? maybe lox-hydrogen starts looking better than lox-methane for ISRU.
So, presumably, if we're not talking about staged combustion, then methane's advantage decreases.
Now *that* is a really solid reason. That's the kind of reason I've been looking for. Methane's mass-specific heat capacity is about three times propane's.So, presumably, if we're not talking about staged combustion, then methane's advantage decreases.
Quote from: strangequark on 02/07/2013 03:42 pmIt'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.Now *that* is a really solid reason. That's the kind of reason I've been looking for. Methane's mass-specific heat capacity is about three times propane's.
Quote from: Proponent on 02/07/2013 04:07 pmQuote from: strangequark on 02/07/2013 03:42 pmIt'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.Now *that* is a really solid reason. That's the kind of reason I've been looking for. Methane's mass-specific heat capacity is about three times propane's.Actually, I now realize I was comparing heat capacities at different temperatures. According to Air Liquide, methane's heat capacity at constant volume is only a bit higher (1714 J/kg/K vs. 1501) at STP.Of course, what's actually relevant are the heat capacities at much higher temperatures. How does methane's high-temperature heat capacity compare with those of other light hydrocarbons?
Even better than purified methane would be commercial LNG (it would save money if the extra purification was not needed).
for either expander cycle or staged combustion, use of O2 for the turbine works better. Because of the greater volume of O2 in any hydrocarbon cycle. See last frame of http://www.microlaunchers.com/7816/L3/sa05/sa05.htmlfor my idea of staged combustion. The O2 temperature can be low--even room temperature to supply enough energy to run a turbine.The Russians have been doing well with the idea for a long time. With LOX side power, it should be fairly easy to change the fuel choice for an engine.
Forked from Jim's Rational SpaceX thread:PM'd you guys but sure let's continue here, interesting subject!Quote from: strangequark on 02/06/2013 07:33 pmQuote from: Robotbeat on 02/06/2013 05:35 pmQuote from: R7 on 02/06/2013 05:34 pmQuote from: Robotbeat on 02/06/2013 05:28 pm(And of course the first one is talking about SSTO with the volume of the craft remaining constant.)Which is the point. Switching to methane does not help, unless you design a bigger vehicle to go with it.No, the conclusions only apply to SSTO and somewhat to first stages (according to your second one).Again, your second link said this:"So Impulse Density doesn’t mean a whole lot for upper stages that don’t interact with the atmosphere, but it does have a little bit of meaning when applied to boosters. For a booster stage, higher impulse density can result in a heavier, but smaller vehicle (generally)."R7, attached is a graph that shows the relative increase in Isp required to counteract the performance decrease due to density decrease. Rho_t is the mass of stage required to enclose a unit volume of propellant. Centaur is about 25, Atlas is around 70, Delta is around 50, Zenit is around 100, as a framework (kg/m^3). I would be happy to discuss assumptions, derivation, and sensitivity later tonight, but we should split the thread if so.If I read the chart correctly a switch from kerosene/RP-1 to methane would (at minimum, 25t stage case?) require about 5% Isp increase to counter the density drop, yes?The dunnspace link I posted listed 347.8s for methane and 338.3 for RP-1, so that's 2.8% increase for methane which is ... not enough?edit: the dunnlink: http://dunnspace.com/alternate_ssto_propellants.htm
The Russians figured out O2 for the turbines enough for a single short flight. Given the damage heated O2 does, would the turbines be easily re-used many times? Perhaps new breakthroughs would be needed.
RD-170 engine for the “Energia” launch-vehicle is intended for reusable operation and is certificated for 10-multiple use. One of the engines was tested at a fire bench up to 20 times.
something like this?QuoteRD-170 engine for the “Energia” launch-vehicle is intended for reusable operation and is certificated for 10-multiple use. One of the engines was tested at a fire bench up to 20 times.source: http://www.npoenergomash.ru/eng/engines/rd171m/
Since this is SpaceX, I would assume that their plan is to run a natural gas pipe, make their own liquid methane, and sort out for themselves what impurities they can tolerate.
(solar) electricity CO2 and H2O in, O2 and CH4 out. Just like on Mars.
Quote from: Lar on 03/27/2013 05:19 pm(solar) electricity CO2 and H2O in, O2 and CH4 out. Just like on Mars. On Earth there's a much simpler route, using bioreactor (fancy name for pretty much any gas/liquid tight container);crap in (really, whatever biomass, municipal waste, poop), biogas out (mostly CH4, CO2). Relatively easy to refine into practically pure LCH4 because the biogas lacks any longer chain hydrocarbons.
I believe we've had posts that say biogas comes with contaminants that would need to be purified out (eg H2S, IIRC).ISTM that the process of refrigerating to a cryo liquid would make it very easy to do some fractional distillation to tidy that up, though.
Question is, where do you get the, erm, feedstock over there near LC-40?
"Thank you for collaborating in project Thunderpants"
Well, I guess methane is in the thread title...Rocket related inspiration doesn't begin until 27 seconds in.
Quote from: go4mars on 03/28/2013 01:34 amWell, I guess methane is in the thread title...Rocket related inspiration doesn't begin until 27 seconds in. In another win for humor, the exhaust plume appears to be from kerolox engines. This kid must've had some serious digestive problems to be producing kerosene vapors.
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 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.
Ethylene may be great once you get it inside the combustion chamber but will it behave properly when driven thru turbopumps and coolant channels? Thermal stability, tendency to polymerize and all that?
What's the route from basic Martian resources to synthesize ethylene?
3. Relative ease of ISRU methane production on Mars compared to heavier hydrocarbons (though this argument does not make make much sense to me); and
That the synthesis of methane is easier than that of heavier hydrocarbons makes perfect sense. What I doubt is that plans for Martian ISRU in the future much influenced SpaceX's choice of propellants for the near term (Raptor).
Try subcooled propylene as a propellant. Even better than ethylene.
Liquefied natural gas enhances affordability and reusabilityLiquefied natural gas is commercially available, affordable, and highly efficient for spaceflight. Unlike other rocket fuels, such as kerosene, liquefied natural gas can be used to pressurize a rocket’s propellant tanks. This is called autogenous pressurization and eliminates the need for costly and complex pressurization systems, like helium. Liquefied natural gas also leaves no soot byproducts as kerosene does, simplifying engine reuse.
You need a catalyst for ethylene to polymerise, so the liquid may still be stable under heat and pressure. According tohttp://www.chemguide.co.uk/mechanisms/freerad/polym.htmlyou need 200 C and 2000 atm of pressure. Typical staged combustion engines are only 200 atm, so this is probably well below what could cause a problem.
Robotbeat, Niloff: what I'm thinking is that lots of people are keen on lox-methane these days, like BO, ULA and Firefly (last I heard Firefly had switched to lox-RP-1, but methane was originally baselined), and they're not all focused on Mars. On the BE-4 info sheet, BO says this about its choice of LNG:Quote from: Blue OriginLiquefied natural gas enhances affordability and reusabilityLiquefied natural gas is commercially available, affordable, and highly efficient for spaceflight. Unlike other rocket fuels, such as kerosene, liquefied natural gas can be used to pressurize a rocket’s propellant tanks. This is called autogenous pressurization and eliminates the need for costly and complex pressurization systems, like helium. Liquefied natural gas also leaves no soot byproducts as kerosene does, simplifying engine reuse..Another factor is that even if SpaceX succeeds wildly, it's going to launch one heckuva lot of mass from the the surface of the Earth before it produces its first liter of ISRU methane. Since there's a lot of hardware to be developed on the way to that first liter of ISRU methane anyway, I would think it would make sense to optimize for Earth launch for the time being and then tweak propulsion systems as needed later for ISRU methane.So, I could believe that if there were several propellant combinations that were approximately equally attractive, SpaceX would probably choose the one that's best for Mars. But if there were other fuels nearly as attractive as methane, why are none of the less-Mars-obsessed players pursuing them?
Very generally on the top of specific impulse and impulse density, I was thinking about optimal mixture ratios. If oxidizer and fuel have different densities, then impulse density will peak at a mixture ratio corresponding to a higher propellant bulk density than where the specific impulse peaks. The larger the difference in the densities of oxidizer and fuel, the larger will tend to be the difference in the mixture ratios of the two peaks. Consider lox-hydrogen. The attached figure shows specific impulse1 and impulse density as a function of both bulk density (lower horizontal axis) mixture ratio (O/F: upper horizontal axis)2.Obviously specific impulse peaks at a density of about 317 kg/m3 (O/F=4., while impulse density peaks at 633 kg/m3 (O/F=17.. Lox-hydrogen stages usually operate at mixture ratios of about 5 or 6. You can imagine situations where you'd want to go significantly higher than that. But you'd never want to go to a bulk density lower than that of maximum specific impulse or higher than that of maximum impulse density. Anyway, there's nothing very profound about this, but I thought I'd mention it, because it hadn't occurred to me before.1. Specifically, these values are 95% of the ideal vacuum values calculated with RPA Lite; the densities of both propellants correspond to those at their respective normal boiling points.2. Note, though, that the mixture-ratio axis is non-linear, because bulk density is not a linear function of mixture ratio (though specific volume, the reciprocal of density, is).
Good point. For perfect burning, mixture ratio should be 16.
It would seem that RS-68 uses too small mixture ratio;
Quote from: hkultala on 04/14/2016 09:01 pmGood point. For perfect burning, mixture ratio should be 16.You mean 8, right?
QuoteIt would seem that RS-68 uses too small mixture ratio;Maybe it's partly to increase the thrust-to-weight ratio.
Hmm, a thought just occurred to me, wherein cryochilled propane's viscosity could be an advantage. Are you familiar with the research on metalized gel propellants? The concept is to add gelling agents like fumed silica to allow you to suspend metal dusts like aluminum powder in the liquid rocket fuel. Aluminum combustion gives off a great deal of energy for its mass, providing additional heat to the exhaust stream. But you know, gelling basically means "increasing the viscosity". With cryochilled propane, you already have increased viscosity vs. "runny" fuels like RP-1 (whether it's sufficient to suspend aluminum particles without gelling agents, that I can't say).
To further develop my understanding of the trade between specific impulse and density, I've done a little thought experiment on ground-launch stages.
1. Except for JP-5 (the composition of which I don't know),
2) Some cooling-related thing makes the engine T/W not scale with propellant density. The extra unburned H2 makes the engine run cooler and allow less mass to be used for cooling?
I've never seen any indication of a 'viscosity' issue with cryo-propane. It's denser than when liquid under normal pressure/temperature but nothing that impedes either turbo-pump or pressure fed use. When tested in the RL10 it was less 'viscos' than RP1 and more like LH2 which was a cited advantage in that type of engine.
So if I read all this right then I get that balancing both impulse and density has not been as straight-forward as even the experts (I'm thinking all the early "when we have hydrogen we can do anything" rocket scientist here ) had thought. Further it would seem that in a TSTO system it might be more efficient to consider different propellants for booster and upper stage despite a slightly higher operations costs?
JP-5 or JP-10? JP-5 Material Safety Data Sheets, (MSDS) are available on-line and composition properties sheets IIRC ....
Quote from: Impaler on 06/09/2015 03:22 amSteven 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.Refrigeration advantages are moot when sharing a thermal environment with LOx. Ethylene is a moderate problem there, because to maintain it in liquid phase at the same temperature you would need to raise LOx tank pressure to 5+ atmospheres (ethylene freezes at the boiling point of LOx at about 3.5atm). That's manageable, but adds weight. Zubrin is currently working on ethylene-N2O green hypergolics - http://www.parabolicarc.com/2015/05/15/pioneer-astronautics/
I've been told on several occasions that Lox-Methane is a very feasible mixture for a Mars mission, with references to the Morpheus lander, though I was curious about its ignition. The advantages I saw listed were non-toxicity, low cost of propellant, and lower energy use versus hypergolic propellants. The obvious disadvantages were impulse density and ignition issues. Aside from gaining hypergolic ignition with an Ethylene-Nitrous Oxide mix, how does its compare with the obvious Lox-Methane and Lox-Ethylene alternatives in terms of advantages and disadvantages?
Quote from: Hyperion5 on 04/16/2016 01:10 amI've been told on several occasions that Lox-Methane is a very feasible mixture for a Mars mission, with references to the Morpheus lander, though I was curious about its ignition. The advantages I saw listed were non-toxicity, low cost of propellant, and lower energy use versus hypergolic propellants. The obvious disadvantages were impulse density and ignition issues. Aside from gaining hypergolic ignition with an Ethylene-Nitrous Oxide mix, how does its compare with the obvious Lox-Methane and Lox-Ethylene alternatives in terms of advantages and disadvantages?Attached are plots showing the performance of both oxygen and nitrous oxide with light hydrocarbons. Low-orbit speed on Mars is about 3.5 km/s. Since the atmosphere is thing and the gravity weak, perhaps 4 km/s is not too optimistic as a delta-V for getting from the surface to low orbit. The first plot shows that both oxygen and nitrous do pretty well. Oxygen is better, but nitrous isn't too bad.On the other hand, the thin atmosphere means that a martian SSTO making a round trip would have a pretty large delta-V to perform in returning to the surface. Taking an approximate, worst case, suppose the total delta-V to orbit and back is 8 km/s. Then, as the second plot shows, the performance of nitrous is really rather poor.
... the thrust should be about linearily propotional to the propellant density ....
I'm interested in seeing plain N2O myself. I always wondered at why it was not used given its ease of storage, self pressurization and ability to be A) a monopropellant and B) a ignition source with a reusable catalyst. It's performance really isn't THAT bad, but in a reusable system that isn't 100% performance optimized I always wondered why it didn't never really saw much consideration.
through my amateurish calculations expect a high 240s-range sea-level ISP
******************************************************************************* NASA-GLENN CHEMICAL EQUILIBRIUM PROGRAM CEA2, MAY 21, 2004 BY BONNIE MCBRIDE AND SANFORD GORDON REFS: NASA RP-1311, PART I, 1994 AND NASA RP-1311, PART II, 1996 ******************************************************************************* prob rocket fac p,bar=200.0 ions pi/pe=1276.851685 mdot=2223.8 reac fuel=N2O moles=1.0 t(k)=300.0 outp short end THEORETICAL ROCKET PERFORMANCE ASSUMING EQUILIBRIUM COMPOSITION DURING EXPANSION FROM FINITE AREA COMBUSTOR Pin = 2900.8 PSIA MDOT/Ac = 2223.800 (KG/S)/M**2 Pinj/Pinf = 1.003364 CASE = REACTANT MOLES ENERGY TEMP KJ/KG-MOL K FUEL N2O 1.0000000 81671.539 300.000 O/F= 0.00000 %FUEL=100.000000 R,EQ.RATIO= 0.000000 PHI,EQ.RATIO= 0.000000 INJECTOR COMB END THROAT EXIT Pinj/P 1.0000 1.0068 1.8334 1276.85 P, BAR 200.00 198.66 109.09 0.15664 T, K 1908.42 1907.02 1670.30 305.99 RHO, KG/CU M 3.6988 1 3.6766 1 2.3049 1 1.8065-1 H, KJ/KG 1855.63 1853.80 1550.33 7.8034 U, KJ/KG 1314.91 1313.48 1077.06 -78.903 G, KJ/KG -12294.6 -12287.8 -10835.9 -2261.27 S, KJ/(KG)(K) 7.4146 7.4156 7.4156 7.4156 M, (1/n) 29.345 29.345 29.344 29.342 (dLV/dLP)t -1.00007 -1.00007 -1.00003 -1.00000 (dLV/dLT)p 0.9998 0.9998 0.9998 1.0000 Cp, KJ/(KG)(K) 1.3049 1.3046 1.2594 0.9960 GAMMAs 1.2771 1.2772 1.2901 1.3976 SON VEL,M/SEC 831.0 830.7 781.4 348.1 MACH NUMBER 0.000 0.073 1.000 5.522 PERFORMANCE PARAMETERS Ae/At 8.0991 1.0000 51.862 CSTAR, M/SEC 1106.7 1106.7 1106.7 CF 0.0547 0.7060 1.7370 Ivac, M/SEC 8993.6 1387.1 1967.5 Isp, M/SEC 60.5 781.4 1922.4 MOLE FRACTIONS *NO 0.00676 0.00673 0.00299 0.00000 NO2 0.00024 0.00024 0.00013 0.00000 *N2 0.66324 0.66325 0.66515 0.66667 *O 0.00001 0.00001 0.00000 0.00000 *O2 0.32974 0.32975 0.33173 0.33333 * THERMODYNAMIC PROPERTIES FITTED TO 20000.K
To further develop my understanding of the trade between specific impulse and density, I've done a little thought experiment on ground-launch stages.In 1996, John Whitehead wrote a cute little paper about SSTO mass budgets (4th attachment to this post. For a few different propellant combinations, he used the rocket equation to calculate the mass ratios needed for a delta-V of 10 km/s (i.e., Earth to LEO with losses). Then he estimated the masses of engines, tanks, pressurants and residual propellants (the last two can be larger than you expect) as a fraction of burn-out mass. Let's call the part of the burn-out mass that's not devoted to those four things the available mass (I'm open to suggestions for a better term). Many subsystems will have to be crammed into this so-called available mass: landing gear, if any, avionics, etc., etc. The idea, though, is that the masses of such subsystems will be approximately independent of the propellants chosen. Hence, the vehicle with the highest available mass fraction should have the largest payload fraction as well.The name of the game, then, is to choose the propellant combination that maximizes the available mass.About the same time as Whitehead's paper, Bruce Dunn presented an analysis in a similar spirit (3rd attachment). Although Dunn's assumptions were perhaps a bit more ad hoc, he covered a wider range of propellants.I've made a similar calculation similar to Whitehead's. There are just two major differences. Firstly, Whitehead assumes that the lift-off thrust-to-weight ratio of the engine is a linear function of propellant density and is 100 for lox/RP-1 and 50 for lox/hydrogen. In contrast, I assume the ratio is proportional to the impulse density of the propellants (it seems to me this makes more sense; any comments?), taking a value of 123 for lox/RP-1 at a typical mixture ratio (essentially, the NK-33 or the AJ-26).The second significant difference is that rather than assuming a particular mixture ratio, I adjust the mixture ratio for maximal performance.Otherwise, to oversimplify slightly, I use pretty much the same assumptions: 10 km/s of delta-V, tanks weigh 10 kg/m3, pressurants and residuals are each 0.25% of the initial propellant load. Specific impulses come from RPA Lite 1.2.8 and are scaled by 0.95 from ideal vacuum values. Chamber pressure is 20 MPa and the area expansion ratio is 40:1. For the time being, propellants are assumed to be at the lower of room temperature and the normal boiling points.Have a look at the first plot attached. It shows specific impulses delivered by various fuels1 burned with oxygen as a function of propellant bulk density. Also shown as grey curves are contours of constant "available mass." These contours are easily calculated, since all that's required in Whitehead's model is a specific impulse and a propellant density. The first table, below, gives optimal figures for each of 30 propellant combinations.Hydrogen does poorly. If the mixture ratio is allowed to vary during flight in an optimal way, the available mass fraction with hydrogen as a fuel increases2 by about 0.026. Other fuels don't benefit much from mixture-ratio variation, so the this enough to boost hydrogen to the middle of the table. But, the substantially larger mass of hydrogen tanks arising from the need to insulate them has been neglected. Taking this into account would knock hydrogen right back to the bottom of the table.Speaking of the table, a couple of columns may not be self-explanatory:* Mix: Linear function of the mixture ratio, being zero for maximum Isp and unity for maximum impulse density.* T/W: Thrust-to-weight ratio of the engine at lift-off (giving the a ratio of 1.3 for the vehicle).* Den exp: slope of the log Isp-log(bulk density) curve at the optimum; shows the relative importance of density compared to Isp.People often obsess about maximizing specific impulse. The Mix column shows that's not generally what you want to do.The "Den exp" column shows the relative sensitivity of available mass fraction to density as opposed to specific impulse. For the better performing propellant combinations, it's about 0.23, meaning that a the figure of merit is approximately: (specific impulse)(bulk density)0.23for an SSTO. This is, of course, somewhat model dependent, but it happens to be about the same as what I estimated from Dunn's results some time ago.OK, so, what about hydrogen peroxide, with its high density? Please have a look at the second plot. This time I've left hydrogen out so as to make the hydrocarbons more visible. As you easily see, peroxide's density does not raise bulk density enough to make up for its lower specific impulse. Bruce Dunn told us that a long time ago, but I find it educational to see it graphically. I also looked at nitric acid, which is even denser (1510 kg/m3) than peroxide (1460 kg/m3). It, however, suffers from lower specific impulse and lower bulk density than you might expect: the fact that it contains quite a bit of free oxygen means that mixture ratios with nitric acid tend to be low.If we consider a delta-V of just 4 km/s -- see the third plot and second table -- peroxide looks much better. As you'll see from the table, the figure of merit at this delta-V, which could correspond to a first stage or a martian SSTO, is something like: (specific impulse)(bulk density)0.4 ,Finally, consider a very low delta-V, like 40 m/s, as shown in the final plot. In this case, impulse density reigns, and peroxide is the run-away winner. The associated table shows that the figure of merit is very close to (specific impulse)(bulk density) ,i.e., impulse density, which is just what you expect when delta-V is small compared to exhaust velocity. Note, though, that we do have to go to very low delta-V's before impulse density dominates.All of the above is applies to ground-lit stages. For upper stages, mass will be more important, since the stage's propellant must be accelerated by lower stages. Hence, the density exponent in the figure of merit will tend to be smaller.1. Except for JP-5 (the composition of which I don't know), the color of each curve is the number of carbon atoms, modulo 10, in each fuel's principal chemical component (e.g., 1 for methane, 2 for ethane and ethylene) expressed in the resistor color code. Solid lines are used for saturated hydrocarbons (alkanes). The two alkenes, ethylene and proplylene, are shown with dashed lines.2. If a different mixture ratio is allowed for each successive 1% of the total propellant volume, the ratio ranges from 17.8 (633 kg/m3) at lift-off to 5.7 (350 kg/m3) at burn-out. A variable mixture-ratio program helps in two ways. Firstly, it simply helps with the rocket equation by allowing more impulse to be packed in at the beginning, where mass doesn't matter so much, while going for higher specific impulse at later times. Secondly, it increases the lift-off thrust-to-weight ratio of the engine, allowing for a smaller engine.EDIT: Added "bulk" to very-low-delta-V figure of merit.
, tanks weigh 10 kg/m3
pressurants and residuals are each 0.25% of the initial propellant load
Chamber pressure is 20 MPa and the area expansion ratio is 40:1.
Hydrogen does poorly. If the mixture ratio is allowed to vary during flight in an optimal way, the available mass fraction with hydrogen as a fuel increases2 by about 0.026.
But, the substantially larger mass of hydrogen tanks arising from the need to insulate them has been neglected.
* Mix: Linear function of the mixture ratio, being zero for maximum Isp and unity for maximum impulse density.* T/W: Thrust-to-weight ratio of the engine at lift-off (giving the a ratio of 1.3 for the vehicle).* Den exp: slope of the log Isp-log(bulk density) curve at the optimum; shows the relative importance of density compared to Isp.
(specific impulse)(bulk density)0.23
Quote, tanks weigh 10 kg/m3Can this concept be defended? Mass loadings on the tanks will certainly be different with different propellants at the very least. Not to mention the x^3/x^2 volume/surface area scaling issue....
...When you're talking about SSTOs, the picture doesn't change. You just can't use the sort of low-ISP high-thrust stage you'd use with a staged rocket. You still have to use H2 because SSTOs are even more ISP-dependent than staged rockets:https://en.wikipedia.org/wiki/File:SSTO_vs_TSTO_for_LEO_Mission.tifYou simply can't get a plausible structural coefficient with a low ISP propellant mix. It just doesn't work.
The volume/surface area scaling issue does not apply to pressure vessels (unless you run into minimum-gauge issues or decide to use a large amount of insulation). Since rocket tanks are fairly well approximated as pressure vessels
If you get really, REALLY tall (like Saturn V first stage size), then you have to start taking into account pressure head (and this can actually allow you to SAVE weight, since you can use a little less ullage pressure and the top of the stage can thus be made a little thinner), but for our purposes here, that's a pretty good estimate.
This last one is perhaps the biggest reason why pump-fed rocket engines are used instead of pressure-fed.
Rei, have you read Dunn's report on various SSTO propellant combinations? It is not kind to hydrogen.http://web.archive.org/web/20120303152352/http://www.dunnspace.com/alternate_ssto_propellants.htmHydrogen may have the best Isp, but liquid hydrogen is, in fact, the least dense liquid known to humankind. It has been worshipped by aerospace since Tsiolkovsky, but in no way is it an optimal fuel for a SSTO rocket, particularly a reusable one (where dry mass is yet more important). Please re-examine your prejudices in light of that Dunn report.
In the current model, most propellant combinations beat hydrogen/oxygen. This is a direct result of assuming a constant-size rather than constant-mass vehicle for all propellants, regardless of density.
Falcon 9 v1.0 was thought to have an ullage pressure of about 50psi, that's more than just "a fraction of an atmosphere overpressure."
Additionally, Saturn V is a poor example because the different stages were built by different entities. Additionally, the first stage is obviously going to be built much different than the other stages due to the lower penalty for high dry mass first stage (with its big ol' fins, etc). You should be comparing pump-fed upper stages to other pump-fed upper stages.
QuoteThe volume/surface area scaling issue does not apply to pressure vessels (unless you run into minimum-gauge issues or decide to use a large amount of insulation). Since rocket tanks are fairly well approximated as pressure vesselsNo, they are not.Even balloon tanks have an overpressure a fraction of an atmosphere. As a general rule, the only high pressure tank in a rocket is a helium pressurant which is steadily released to ensure a stable supply of propellant to the turbopumps.You simply cannot multiply volume by a constant. That's not at all an accurate representation of tankage mass. Example: Saturn V first stage = 130 tonnes, holds 1305 cubic meters of propellant = 100kg/m^3. Second stage = 38 tonnes, 1559 cubic meters = 24kg/m^3. Third stage = 10 tonnes, 326 cubic meters = 31kg/m^3. Not. Even. Close.You simply cannot take some sort of linear scaling parameter with volume to estimate the tankage mass. Rockets just don't work that way. Using the posted formula above one would come to the conclusion that Saturn V's hydrogen stages' would be *four times heavier* than they actually were.QuoteIf you get really, REALLY tall (like Saturn V first stage size), then you have to start taking into account pressure head (and this can actually allow you to SAVE weight, since you can use a little less ullage pressure and the top of the stage can thus be made a little thinner), but for our purposes here, that's a pretty good estimate.And as was just demonstrated above, the Saturn-V first stage is half of an order of magnitude heavier per unit volume, not lighter. QuoteThis last one is perhaps the biggest reason why pump-fed rocket engines are used instead of pressure-fed.Which is also why they're not pressure vessels. Except in balloon tanks internal pressure is only kept high enough to keep the turbos fed. And even with balloon tanks, it's hardly something one would consider a "pressure vessel".QuoteRei, have you read Dunn's report on various SSTO propellant combinations? It is not kind to hydrogen.http://web.archive.org/web/20120303152352/http://www.dunnspace.com/alternate_ssto_propellants.htmHydrogen may have the best Isp, but liquid hydrogen is, in fact, the least dense liquid known to humankind. It has been worshipped by aerospace since Tsiolkovsky, but in no way is it an optimal fuel for a SSTO rocket, particularly a reusable one (where dry mass is yet more important). Please re-examine your prejudices in light of that Dunn report.All serious efforts toward SSTOs have used hydrogen. There is a reason for this. And that reason is in the graph that I posted.You want to prove NASA wrong? Start with at least posting something that's been peer-reviewed. Even with just a cursory glance I can see glaring problems with the stated work, such as how he constrains all systems to have the same propellant volume. But of course, the author is kind enough to mention this:QuoteIn the current model, most propellant combinations beat hydrogen/oxygen. This is a direct result of assuming a constant-size rather than constant-mass vehicle for all propellants, regardless of density.
The use of a common bulkhead saved 3.6 tonnes in weight, both by eliminating one bulkhead and by reducing the overall length of the stage.
Quote from: Robotbeat on 04/27/2016 01:59 amFalcon 9 v1.0 was thought to have an ullage pressure of about 50psi, that's more than just "a fraction of an atmosphere overpressure."1) Seriously, why do people on this site use arcane measurements like psi, mmHg, etc? Use metric people, this isn't the dark ages....2) Do you have a solid reference for that? Atlas, the classic example of a balloon tank, was 34kPa:https://en.wikipedia.org/wiki/SM-65_AtlasFalcon 9 is partially pressure stabilized, but not to the degree of Atlas (Atlas couldn't even be transported or left on the stand unpressurized)But let's just say that it's 350 kPa. Shortly after launch it's pulling 2G. ...
Now can you try subchilling the propellants like Dunn?
...
Assuming a perfect burn, no frozen combustion?
QuoteHydrogen does poorly. If the mixture ratio is allowed to vary during flight in an optimal way, the available mass fraction with hydrogen as a fuel increases2 by about 0.026.I'm confused by this argument. On what grounds are you determining "poor performance" and "optimal"? Maybe I missed something, because your graph rightfully shows hydrogen's ISP vastly superior to the others (but its bulk density, obviously, vastly lower - no shockers there)
QuoteBut, the substantially larger mass of hydrogen tanks arising from the need to insulate them has been neglected.Depends on the context, of course. On Earth, uninsulated H2 tanks liquefy the surrounding air, causing an extremely rapid heat loss. On Mars, however, not only is it easier to maintain a much lower radiative equlilibrium, and not only is there far, far less convective losses, but the air doesn't liquefy; like water vapour on LOX tanks, it freezes at LH temperatures, providing an insulative ice that falls off during launch. It's not immediately obvious that LH tanks for rockets on Mars would require significant, if any, insulation (obviously long-term storage needs insulation)
Quote* Mix: Linear function of the mixture ratio, being zero for maximum Isp and unity for maximum impulse density.* T/W: Thrust-to-weight ratio of the engine at lift-off (giving the a ratio of 1.3 for the vehicle).* Den exp: slope of the log Isp-log(bulk density) curve at the optimum; shows the relative importance of density compared to Isp.I don't see these things in your graphs.
Quote (specific impulse)(bulk density)0.23I can see no justification for the usage of a formula involving a linear multiplication of specific impulse and bulk density.
There's no real mystery here about the optimums for chemical rockets with current propellants. This has been worked out long, long ago. Hydrocarbon first stages provide massive thrust (due to the high propellant density) and the tankage costs are kept low (due to the high density and simpler construction) on the massive first-stage tanks. Hydrogen provides the high ISP needed for subsequent stages to keep the size requirements on the first stage down.
Quote from: Rei on 04/26/2016 10:46 pm...By way of adding to what Robotbeat and dkovacic have said, let me suggest that you read the Whitehead paper I mentioned previously. Whitehead went to some lengths to justify the scaling laws he used.
QuoteAssuming a perfect burn, no frozen combustion?The specific-impulse figures assume shifting equilibrium.
You're actually using precisely the same metric when you assert that lox-hydrogen is the superior SSTO propellant combination because of its specific impulse. It's just that you're just asserting that the exponent is zero.
Please note that I'm not claiming to have discovered anything new. As stated in the very first sentence of my post, my goal is simply "To further develop my understanding of the trade between specific impulse and density." As to upper stages, I write "All of the above is applies to ground-lit stages. For upper stages, mass will be more important, since the stage's propellant must be accelerated by lower stages. Hence, the density exponent in the figure of merit will tend to be smaller." That's something I want to look into in a little more detail.
By the way, there are a couple of factors working against hydrogen that are not factored into my simple analysis (or into Whitehead's or Dunn's). First of all, its bulkiness raises drag losses. More subtly, its high specific impulse reduces acceleration and increases gravity losses.
Ack, what happened to my earlier post? I had written a long post full of something like 20 different rockets with their dry masses and propellant volumes... and it's since disappeared. Argh I don't have the time to write it all again.... from memory it was something like: LOX/LH: Delta-IV heavy, Ariane V core, Ariane V upper, Energia core, Centaur, and several others were compared. With the exception of the Ariane V upper, they were in around the mid 30s kg/m^3. I remember that Ariane V core was best, at 28kg/m³LOX/RP1: Energia booster, Atlas V, Falcon 9, Soyuz FG, and a variety of other rockets were compared. Most were in the ~75-120 range, but Falcon was lower, I think around 60 kg/m^3 (SpaceX really does impressive work)First stages tended to be a heavier than second, but thirds tended to also be a little heavier than seconds**. Boosters tended to be heavier than firsts. Size seemed to play a role, but "technology level" seemed to play a bigger one. The biggest factor however was propellant combination; LOX/LH clearly gets a far lower kg/m³ than LOX/RP1 in whole, real-world stages.** - I would wager that the reasons are:1) Firsts have to fight gravity losses, so need heavier thrusters. They also have to bear more weight.2) Second and higher get vacuum ISPs. This comes at the cost of heavier nozzles, but obviously it's worth the mass.3) There appears to be (to some degree, for at least over some range) "economies of scale" in rocketry masses, which third stages suffer from relative to second. They also sometimes face additional hardware needs (such as RCS, restartable engines, etc) that their lower stages don't.The basic takeaway is that you absolutely cannot pretend that there's some simple linear, propellent-independent relationship between tank volume and tank mass.
All serious efforts toward SSTOs have used hydrogen.
You want to prove NASA wrong? Start with at least posting something that's been peer-reviewed. Even with just a cursory glance I can see glaring problems with the stated work, such as how he constrains all systems to have the same propellant volume. But of course, the author is kind enough to mention this:QuoteIn the current model, most propellant combinations beat hydrogen/oxygen. This is a direct result of assuming a constant-size rather than constant-mass vehicle for all propellants, regardless of density.
Atlas, the classic example of a balloon tank, was 34kPa:https://en.wikipedia.org/wiki/SM-65_Atlas
And all of this is irrelevant anyway, because tanks demonstrably do not have any sort of linear, propellant-ambivalent correspondence between volume and dry mass. Look up tank masses. It just doesn't work that way.
So I would say it brings it closer, but I would agree that for real world examples, dry mass does tend to be significantly higher for RP-1/LOX. Based on comparison of Ariane V and Atlas V, I would say that ratio of 2:1 is realistic. But why is that so is not obvious to me.
Arguably the most serious SSTO effort ever was Roton, in that it got as far as very early flight tests of an actual SSTO design as opposed to a technology demonstrator. Roton burned lox and RP-1.
There's a difference between stage mass and tank mass, as Whitehead discusses.
You'll see that once the mass of the "propellant module" exceeds 1000 lbm, propellant mass is quite proportional to module mass.
It's also the case that propellant mass is inversely proportional to to tank pressure, which is consistent with pressure-vessel principles.
As for the link you provided: it's not a simple fuel substitution, the author (again, not peer reviewed) is changing all sorts of other parameters as well such as the wing design*. It's also "proof by ghost reference":Using the same methodology for calculating massesThat'd be nice if we actually had the methodology onhand. If it were peer reviewed I could at least say "well, somebody's checked his numbers". Out of context, their numbers make no sense - for example, the fuel tank mass is listed as "-". What does "-" mean? That it has no fuel tank?
Quote from: Rei on 04/27/2016 04:43 pmAs for the link you provided: it's not a simple fuel substitution, the author (again, not peer reviewed) is changing all sorts of other parameters as well such as the wing design*. It's also "proof by ghost reference":Using the same methodology for calculating massesThat'd be nice if we actually had the methodology onhand. If it were peer reviewed I could at least say "well, somebody's checked his numbers". Out of context, their numbers make no sense - for example, the fuel tank mass is listed as "-". What does "-" mean? That it has no fuel tank?You missed the reference I take it? "The NASA Access to Space LO2/hydrogen single stage to orbit rocket was examined, and the configuration reaccomplished with LO2/kerosene as the propellants."For reference since you didn't look it's here, (http://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/19940022648.pdf) page 40 onwards.
....All serious efforts toward SSTOs have used hydrogen.....
Quote from: dkovacic on 04/27/2016 02:54 pmSo I would say it brings it closer, but I would agree that for real world examples, dry mass does tend to be significantly higher for RP-1/LOX. Based on comparison of Ariane V and Atlas V, I would say that ratio of 2:1 is realistic. But why is that so is not obvious to me.One major factor seems fairly obvious to me - stress at the bottom of the tank due to G forces on the propellant (at rest or in flight) are proportional to propellant mass, not propellant volume.QuoteArguably the most serious SSTO effort ever was Roton, in that it got as far as very early flight tests of an actual SSTO design as opposed to a technology demonstrator. Roton burned lox and RP-1.Wikipedia states (albeit without citations):Rotary Rocket failed due to lack of funding, but some have suggested that the design itself was inherently flawed.The Rotary Rocket did fly three test flights and a composite propellant tank survived a full test program, however these tests revealed problems. For instance, the ATV demonstrated that landing the Rotary Rocket was tricky, even dangerous. Test pilots have a rating system, the Cooper-Harper rating scale, for vehicles between 1 and 10 that relates to difficulty to pilot. The Roton ATV scored a 10 — the vehicle simulator was found to be almost unflyable by anyone except the Rotary test pilots, and even then there were short periods where the vehicle was out of control.Other aspects of the flight plan remained unproven and it is unknown whether Roton could have developed sufficient performance to reach orbit with a single stage, and return – although on paper this might have been possible. These doubts led some of the aerospace community to dismiss the Rotary Rocket concept as a pipe dream. Whether the concept would have worked successfully remains open to speculation.I tend to agree. What they built had no rocket engine, no heat shield, and was made of a bunch of scrap parts. It was funded in large part by Tom Clancy. It's pretty hard to take seriously. It's like using Armadillo Aerospace as a reference....edit...
LOX Prop Load: 315,364 lbmJP4 Prop Load: 83,000 lbm
Now to the tank mass. As other have pointed out, across the range of interesting and useful propellant loadings, tank mass scales as a function of volume and ullage pressure; nothing else is of any consequence.
LOX Prop Load: 315,364 lbmJP4 Prop Load: 83,000 lbmVehicle Landed Mass: 22,446.7 lbm (as built and calculated from drawings)
Quote from: HMXHMX on 04/27/2016 10:36 pmLOX Prop Load: 315,364 lbmJP4 Prop Load: 83,000 lbmSince this thread is all about the amateur's obsession with propellant choices, could I ask you to expand on the factors that tipped the scales in favor of JP4 in this application? And, if I may try to sound a bit more professional, what was the pressurization system?
$30M would have been a small fundraising round for a new motorcycle company, let alone a rocket company. Getting a handful of angel investors like Clancy is a far cry from actually having enough industry respect to get proper funding. As mentioned previously, it's pretty hard to take seriously as a reference a rocket company as an example whose test rocket wasn't really a rocket and who had to use used parts from a crashed helicopter (of a model that hadn't even been produced in three decades), sold at 1/20th their normal cost because they were salvage, in order to get a test craft off the ground.... a craft that turned out to be exceedingly difficult to actually fly to boot.QuoteNow to the tank mass. As other have pointed out, across the range of interesting and useful propellant loadings, tank mass scales as a function of volume and ullage pressure; nothing else is of any consequence.As has been repeatedly pointed out by the actual stats of actual real-world vehicles, this is demonstrably false. Hydrogen stages generally are upper 20s to upper 30s kg/m³. Demonstrably. Hydrocarbon stages are generally 70-120 kg/m³ (with the exception of SpaceX, which is still far more than H2 stages). DemonstrablyAsserting that things scale otherwise than they do in reality is to deny reality. It doesn't matter how you think things should be - what matters is how they demonstrably actually are. Tanks and stages scale in far closer correspondence to propellant mass than propellant volume.(Nor does the argument that "these are whole stages, not just tanks" improve the case any.)QuoteLOX Prop Load: 315,364 lbmJP4 Prop Load: 83,000 lbmVehicle Landed Mass: 22,446.7 lbm (as built and calculated from drawings)Converting to meaningful numbers, 143047kg of LOX and 37648kg of JP4 with a landed mass of 10181kg. LOX density = 1.141g/cc, JP4 density @15C = .804g/cc. Thus the volumes are 125,4m³ and 46,8m³, respectively, for a total of 172,2m³. Meaning 60kg/m³. Meaning, double the figure for hydrogen.Were you trying to demonstrate that masses do not simply scale proportional to volume? Because that's what you just did.Furthermore, given that the ATV was powered by H2O2, I'm assuming that you're talking about your never-built full scale design. In short, these numbers are purely hypothetical. Versus the actual numbers from actual stages that I've been reporting here.