Author Topic: RP-1, methane, impulse density Q&A  (Read 94766 times)

Offline R7

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RP-1, methane, impulse density Q&A
« on: 02/06/2013 08:12 pm »
Forked from Jim's Rational SpaceX thread:

PM'd you guys but sure let's continue here, interesting subject!

(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
« Last Edit: 04/28/2016 03:09 pm by Chris Bergin »
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Offline Robotbeat

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Re: RP-1, methane, impulse density
« Reply #1 on: 02/06/2013 08:20 pm »
What are the assumptions and equations behind the graph?
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Offline R7

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Re: RP-1, methane, impulse density
« Reply #2 on: 02/06/2013 08:43 pm »
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.
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Offline Lar

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Re: RP-1, methane, impulse density
« Reply #3 on: 02/06/2013 09:31 pm »
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.
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)
"I think it would be great to be born on Earth and to die on Mars. Just hopefully not at the point of impact." -Elon Musk
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Offline Robotbeat

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Re: RP-1, methane, impulse density
« Reply #4 on: 02/06/2013 09:46 pm »
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.
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)
Well, you can split up the methane to hydrogen and carbon monoxide and use that to produce whatever hydrocarbon you want. 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.
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Offline R7

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Re: RP-1, methane, impulse density
« Reply #5 on: 02/06/2013 09:59 pm »
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.

Plus one. The issue with methane as feedstock is yield, processes exist but currently not economically viable. OTOH Mars colony would put economics upside down.

Btw found great paper on effects of impulse density http://www.dtic.mil/dtic/tr/fulltext/u2/283940.pdf, gotta read this with time.

While quickly hopping through it had a revelation why it's more complicated with multiple stages, individual stages take performance hit if volume constrained but upper stages get lighter.
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Offline strangequark

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Re: RP-1, methane, impulse density
« Reply #6 on: 02/07/2013 12:01 am »
What are the assumptions and equations behind the graph?

PDF attached. The previous is a plot of "f" as defined in the attachment. Zero payload, and basically assuming neglible engine mass. If you don't, there's extra constants, so we lose the scale-lessness, but it drives the curves down and closer together regardless.
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.

There are assumptions here, granted, but realize that the density impulse parameter is essentially using the ln(R)=1-R assumption, which is only good for very small delta V (mass ratio close to 1). It's great for evaluating RCS systems where the propellant mass is a small percentage of the vehicle, not so much for full stages.
« Last Edit: 02/07/2013 12:19 am by strangequark »

Offline Proponent

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Re: RP-1, methane, impulse density
« Reply #7 on: 02/07/2013 09:21 am »
[From the previous thread.]

Methane 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.

Offline luksol

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Re: RP-1, methane, impulse density
« Reply #8 on: 02/07/2013 09:45 am »
[From the previous thread.]

Methane 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.

Offline Proponent

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Re: RP-1, methane, impulse density
« Reply #9 on: 02/07/2013 10:52 am »
The 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, and

2.  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.

Offline R7

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Re: RP-1, methane, impulse density
« Reply #10 on: 02/07/2013 11:14 am »
First, I fully realize now that have been way too much speed blinded by the Dunn paper to not see how situation is more complex with multiple stages! Not at all as open-and-shut case as for SSTOs. Shame on me, big thanks especially to Robotbeat and strangequark, and hey, excellent learning experience and good thread ensuing?

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.

I used numbers from the Dunn paper, it assumed fixed pc of 10Mpa and area ratio 100 for all propellants. Isp figures are 90% of theoretical shifting equilibrium.

Impressive improvement with RD-180 for sure, but that kind of defeats the original assumption (by me) that rest of the vehicle stays essentially the same while just replacing old volume of kerolox  with equal volume of methane/LOX (methlox?), and that would be all that it takes to 'help' which I presume meaning increased payload or delta-v. (yeah, that sentence might benefit from splitting...) Dunn pressure is nearly equal to Merlin 1-D so assumed the numbers are valid in this case.

Btw is how does thrust of methane RD-180 compare to original of exact equal size? Gut feeling the mass flux drops slightly (Note to self: get the constitution to cough up the $200 for RPA tool license...)

Changing Merlin to drink methane of course means considerable redesign anyway. Wondering what would be minimal change requirement? Methane as TVC hydraulic fluid?? Just some orifice re-tuning or major TPA redesign?? What a can of worms, gotta pick up Huzel&Huang again  :P

Btw2 one could of course settle the original claim and possible incorrectness of my 'incorrection' if some sort of mass and propellant volume breakdown of F9R is available? Are there numbers for F9? Not much in wiki or spacex.com. But still would have to come to an agreement what assumptions be made in engine department.

Btw3 great pdf!



Btw4 the originating thread briefly mentioned subcooling and of course Dunn paper is all about it's benefits. Gotta ask, if you'd use subcooled LOX/methane/propane etc how is it supposed to stay subcooled in the rocket's tanks while being fueled and waiting for launch? Ordinarily the cryogenics just boil off in manageable rate in ambient (?) pressure so the propellant temperature remains fixed (and density too), yes? But then you pump subcooled stuff into tanks, it starts to warm up and expand. Can't see any margins for some sort of active cooling on the rocket, would there be provisions to recirculate propellant back to GSE for recooling?

In any case, sounds practically challenging. This might lead to some random aerospace company release a carefully worded press release that goes something like

"During final fifteen minute countdown a temporary ground support equipment anomaly caused a rapid volume increase event inside the launch vehicle. Tank walls that are designed to relieve pressure ruptured to protect the vehicle"

Btw5 this post has way too many btws!
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Offline Proponent

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Re: RP-1, methane, impulse density
« Reply #11 on: 02/07/2013 12:03 pm »
Btw found great paper on effects of impulse density http://www.dtic.mil/dtic/tr/fulltext/u2/283940.pdf, gotta read this with time.

Wow, that is a nice paper, so clearly written and with a top-notch abstract up front.  They don't write 'em like that anymore.

My quick scan thus far suggests there may be a couple of issues it doesn't cover, namely the higher gravity losses suffered by high-Isp propellants and the higher drag losses suffered by low-density propellants.

Offline luksol

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Re: RP-1, methane, impulse density
« Reply #12 on: 02/07/2013 12:04 pm »
The 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.

Didn't know that :) I'll have a read.

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.

True. It was sort of my assumption.

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, and

2.  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.

Agreed. It depends if benefits of methane are greater benefits of other propelant(s).

Also, russians have developed a RD-0146U engine that was able to use both methane and hydrogen (though not at the same time ;) ). A single engine was fired several times with hydrogen and methane as a fuel.

Offline R7

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Re: RP-1, methane, impulse density
« Reply #13 on: 02/07/2013 12:11 pm »
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.
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Offline R7

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Re: RP-1, methane, impulse density
« Reply #14 on: 02/07/2013 12:35 pm »
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.

Just found: http://ia700505.us.archive.org/2/items/nasa_techdoc_19850004010/19850004010.pdf

"No significant coking detected" in the last page. Gotta read more carefully. Cerebral digestion tract in pain with all these pdfs  ;D

edit: the related page with description

http://archive.org/details/nasa_techdoc_19850004010

Quote
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.
« Last Edit: 02/07/2013 12:38 pm by R7 »
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Offline Lar

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Re: RP-1, methane, impulse density
« Reply #15 on: 02/07/2013 02:48 pm »
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.
My thinking, admittedly based on a possibly flawed understanding was that water, while likely present, is not going to be hugely abundant on Mars. Certainly not compared to CO2

ISRU propellant manufacture, it seems to me, in essence consists of rearranging atoms from lower energy compounds into higher energy compounds using processes and endothermic reactions that are convenient, storing inputted energy (solar or nuclear)  over a fairly long time... which is then used up in a very short period of time (during burns).

For that reason, it seems to me that LOX-Hydrogen is less desirable on Mars, even if the ISP is higher and the engine more well understood, since it uses a lot of atoms that are relatively scarce for a given amount of deltaV generated.

Methane, since it uses a fair bit of carbon, easy to come by, is more desirable. And of course, carbon monoxide, using no hydrogen at all, seems ideal. Mars has far more carbon than it does hydrogen.

(On the moon the situation is reversed and it's likely that hydrogen will be more plentiful)
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Re: RP-1, methane, impulse density
« Reply #16 on: 02/07/2013 03:42 pm »
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.

Offline Proponent

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Re: RP-1, methane, impulse density
« Reply #17 on: 02/07/2013 04:07 pm »

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.

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.
« Last Edit: 02/07/2013 06:25 pm by Proponent »

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Re: RP-1, methane, impulse density
« Reply #18 on: 02/07/2013 04:15 pm »
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.

A vague memory says that outfit moved to the US.

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Re: RP-1, methane, impulse density
« Reply #19 on: 02/07/2013 04:51 pm »
, what happened to C&Space? IIRC South Korean company

A vague memory says that outfit moved to the US.

Correct memory, found it! The new website is at http://www.darmatechnology.com/chase-10-methane-rocket-engine.html.

The old website (candspace.com) sells health food and ... bags now  ;D


Question on methane's thermodynamics: Is it suitable for expander cycle? High specific heat would imply ... yes?
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Offline go4mars

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Re: RP-1, methane, impulse density
« Reply #20 on: 02/07/2013 05:32 pm »
LNG is cheap and abundant, but pure methane is another matter.
No it isn't.  If there is lot's of ethane, propane, condensates, in the feedstock, midstream facilities can strip them to more than 99% which is commonly done when the price of ethane gets a lot higher than methane.  Many wells in Texas (and elsewhere) produce nearly pure methane and CO2 mixtures anyways (which simplifies things).

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.
1)  The local cost of methane in Texas is a lot lower than LNG. 
2)  For current conditions, that is true.  But if you are aiming toward a gas 'n' go reusable BFR to take tens of thousands to space, and the whole engine and testing program is going to require X amount of fuel, then it is well worth minimizing fuel cost (arguably). 

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.
Depends on architecture choices.  For example, if the upper stage is required to land, then refill on the Martian surface for launch (either for moving stuff around Mars, to Mars orbit, or back to Earth).  Even if a different architecture is chosen to take tourists from Mars to Earth, not needing a separate engine development program (which would be a less-used, perhaps less reliable engine) would save time and money. 

Methane, with its simple C-H bonds, probably is less subject to coking...
Not less coking; No coking.

the tonnage launched from Earth will exceed the tonnage launched from Mars by orders of magnitude,
Orders of magnitude?  Maybe I don't understand what architecture you have in mind.

2.  Setting up ISRU on Mars will require a huge amount of R&D
I keep seeing that assumption.  Take energy, water, CO2, make methane and oxygen.  Engineering will need to be done (whether that is drilling a well to subsurface water, microwaving it out of the ground, moving ice around, etc.) but I don't understanding how this would be a large portion of architecture cost. 

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.
Maybe.  But if presumably you are landing with a methane-engine crasher stage, perhaps bringing your own hydrogen for the first few trips, it seems easier to reuse the same rocket to fly around (or home) rather than bringing a second rocket that needs different insulation, pluming, tanks...   

So, presumably, if we're not talking about staged combustion, then methane's advantage decreases.
True, but not by a lot if you believe price and ease of manufacture&storage on Mars are important factors (which you might not). 
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Offline strangequark

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Re: RP-1, methane, impulse density
« Reply #21 on: 02/07/2013 05:49 pm »
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.

Well, it's good for an expander cycle too. Gas generators let you have a high turbine pressure ratio, so the advantage is not as prominent, and the higher density of the other propellants might be more important.

Offline Proponent

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Re: RP-1, methane, impulse density
« Reply #22 on: 02/07/2013 06:23 pm »

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.

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?
« Last Edit: 02/07/2013 06:24 pm by Proponent »

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Re: RP-1, methane, impulse density
« Reply #23 on: 02/07/2013 07:06 pm »

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.

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?

It's 7 kJ/kg-K at 1000K and 6000psi, which would represent fuel rich drive gas for a staged engine (is actually a bit lower due to oxygen content and reaction products).

2.6 kJ/kg-K at 400K and 600psi, which would be an expander.

Propane is 5 kJ/kg-K at 1000K and 6000psi, and 2 kJ/kg-K for 400K and 600psi. Turbine power is linear with Cp, by the way.

Edit: Added the kelvins
« Last Edit: 02/08/2013 04:48 pm by strangequark »

Offline tnphysics

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Re: RP-1, methane, impulse density
« Reply #24 on: 02/10/2013 08:58 pm »
Even better than purified methane would be commercial LNG (it would save money if the extra purification was not needed).

Offline cpooley

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Re: RP-1, methane, impulse density
« Reply #25 on: 03/25/2013 11:34 pm »
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.html
for 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.

Offline go4mars

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Re: RP-1, methane, impulse density
« Reply #26 on: 03/26/2013 02:44 am »
Even better than purified methane would be commercial LNG (it would save money if the extra purification was not needed).
No. 
LNG generally has a higher concentration of larger molecules than sales nat gas in North America.  Particularly in Texas (where mid-stream companies can get noteably higher prices for stripping out the ethane+ and selling it to chemical companies and refineries). 

Plus, several productive formations within Texas (I'm not willing to research tonight where relative to Boca Chica) will have very high thermal maturation in certain areas, where the natural gas consists almost entirely of CH4 and CO2.
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Offline solartear

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Re: RP-1, methane, impulse density
« Reply #27 on: 03/27/2013 05:35 am »
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.html
for 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.

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.

Offline Hyperion5

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Re: RP-1, methane, impulse density
« Reply #28 on: 03/27/2013 05:48 am »
Forked from Jim's Rational SpaceX thread:

PM'd you guys but sure let's continue here, interesting subject!

(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

Baldusi's been talking up the performance of the RD-162 engine for awhile now, so I thought I'd put up the comparison for it. 

RD-162RD-191Dif
SL Thrust (kN)20001920.84.12%
Dry Weight21002200-4.55%
T/W97899.08%
Chamber Pressure (kg/cm²)175262.6-33%
Chamber Pressure (psi)2,4893,735
SL isp (s)3213113.22%
Vac isp (s)3563385.33%

To me it looks like the RD-162 engine cannot quite make up for its inferior bulk density against an RD-191, but is that quite true?  It also features a bit more thrust, 33% lower chamber pressure, and a better t/w ratio.  Would these factors in addition to its Isp advantage be enough to make the RD-162 as good as booster engine as an RD-191? 

Offline aga

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Re: RP-1, methane, impulse density
« Reply #29 on: 03/27/2013 07:04 am »
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.

something like this?
Quote
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.
source: http://www.npoenergomash.ru/eng/engines/rd171m/
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Re: RP-1, methane, impulse density
« Reply #30 on: 03/27/2013 07:45 am »
something like this?
Quote
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.
source: http://www.npoenergomash.ru/eng/engines/rd171m/

Good point. 20 times does seem like a good start.
« Last Edit: 03/27/2013 07:48 am by solartear »

Offline Lar

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Re: RP-1, methane, impulse density
« Reply #31 on: 03/27/2013 05:19 pm »

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.

Since this is SpaceX, they may well do something similar to this

http://blog.cafefoundation.org/?p=7540

(solar) electricity CO2 and H2O in, O2 and CH4 out. Just like on Mars. :)

Probably not, but it's about as vertically integrated as you can get.

Sorry for replying to an older post... but someone posted a link to this blog and it tickled my memory about this thread.
"I think it would be great to be born on Earth and to die on Mars. Just hopefully not at the point of impact." -Elon Musk
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Offline R7

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Re: RP-1, methane, impulse density
« Reply #32 on: 03/27/2013 06:11 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.
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Offline Lar

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Re: RP-1, methane, impulse density
« Reply #33 on: 03/27/2013 06:45 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.
Right, except that (if I understand these guys, their website is a bit sketchy) they aren't using any crap. (or any other biomass for that matter :) ), just CO2 and water and electricity. But there seem to be some bioorganisms involved so I'm confused.

Anyway, refining pipeline gas to remove everything but the CH4 may be a more feasible approach on earth.
« Last Edit: 03/27/2013 06:45 pm by Lar »
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Offline MP99

Re: RP-1, methane, impulse density
« Reply #34 on: 03/27/2013 06:52 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?

cheers, Martin

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Re: RP-1, methane, impulse density
« Reply #35 on: 03/27/2013 07:51 pm »
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.
CO2, H2S, N2, H2 .. and that's about it. Landfill gas contains more extra stuff. Water scrubbing and regenerable filters get the CO2 and H2S. Distillation the rest.

Quote
Question is, where do you get the, erm, feedstock over there near LC-40?
There are thousands of people working in the Cape on daily basis, no? Get permission to tap the bioreactor into the sewage system. Then even random visitors can "directly participate" in spaceflight. Stickers in the toilets saying "Thank you for collaborating in project Thunderpants" ;)
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Offline Lar

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Re: RP-1, methane, impulse density
« Reply #36 on: 03/27/2013 07:59 pm »
"Thank you for collaborating in project Thunderpants" ;)


I do believe I have my next signature line... Oh, and you owe me a keyboard :) Well, not really.
"I think it would be great to be born on Earth and to die on Mars. Just hopefully not at the point of impact." -Elon Musk
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Offline go4mars

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Re: RP-1, methane, impulse density
« Reply #37 on: 03/28/2013 01:34 am »
Well, I guess methane is in the thread title...
Rocket related inspiration doesn't begin until 27 seconds in. 

« Last Edit: 03/28/2013 01:41 am by go4mars »
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Offline Hyperion5

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Re: RP-1, methane, impulse density
« Reply #38 on: 03/28/2013 03:08 am »
Well, 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.   :D ;D

Offline go4mars

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Re: RP-1, methane, impulse density
« Reply #39 on: 02/13/2014 04:00 am »
Well, 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.   :D ;D
Why do you assume kerolox?  Flame color? 
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Offline Proponent

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Re: RP-1, methane, impulse density
« Reply #40 on: 10/15/2014 10:25 am »
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.

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 Chemistry 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.

EDIT:  Added missing 'r' in "NIST Chemistry WebBook"
« Last Edit: 01/05/2015 11:31 am by Proponent »

Offline Steven Pietrobon

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Re: RP-1, methane, impulse density
« Reply #41 on: 10/24/2014 07:37 am »
Impulse density is a subject I've been interested in for quite a while. Here's a table showing the impulse density of various propellant combinations. The best is F2/NH3! The worst is O2/H2, with O2/CH4 at the low end compared to many other propellants.

MR = Oxidiser to fuel mixture ratio
dp = Propellant density (kilograms per litre)
ve = Effective exhaust speed (divide by g = 9.80665 m/s^2 to get Isp in seconds)
Id = Impulse density (Newton seconds per litre)

Propellants  MR   dp (kg/L)  ve (m/s) Id (Ns/L)
O2/H2        5.0   0.3251     4455     1448
O2/H2        6.0   0.3622     4444     1610
O2/H2        7.5   0.4120     4365     1798
O2/B2H6      2.0   0.7447     4056     3021
O2/NH3       1.4   0.8896     3399     3024
O2/CH4       3.6   0.8376     3656     3062
O2/CH4O      1.4   0.9640     3238     3121
O2/Atsetam   1.8   0.9041     3622     3275
O2/C2H6O     1.9   0.9928     3307     3283   
O2/C2H6      3.2   0.9252     3634     3362
O2/C3H8      3.1   0.9304     3613     3362
O2/C3H4      2.4   0.9666     3696     3573
O2/C3H6      2.7   0.9782     3681     3601
O2/RP–1      2.8   1.0307     3554     3663
O2/C3H4O     1.5   1.0572     3572     3776
O2/C10H16    2.6   1.0471     3608     3778
O2/C7H8      2.4   1.0954     3628     3974

N2O4/B2H6    1.3   0.7196     3859     2777
N2O4/NH3     2.0   1.0428     3097     3230
N2O4/UDMH    2.9   1.1823     3350     3961
N2O4/AZ50    2.2   1.1979     3366     4032
N2O4/MMH     2.4   1.2051     3366     4056
N2O4/N2H4    1.4   1.2156     3371     4097

N2O/C2H6O    5.7   1.1301     3042     3438
N2O/C2H6    10.1   1.1123     3117     3467
N2O/RP-1     9.2   1.1626     3099     3603

HNO3/RP-1    5.2   1.3162     3085     4060

HTP/H2      17.0   0.6925     3592     2487
HTP/B2H6     1.84  0.7947     3970     3156
HTP/NH3      3.1   1.1216     3068     3441
HTP/CH4      8.5   1.1440     3245     3712
HTP/C2H6     8.0   1.2245     3248     3978
HTP/C3H8     7.8   1.2284     3242     3982
HTP/C4H6     6.9   1.2274     3274     4019
HTP/C3H4     6.6   1.2573     3319     4173
HTP/C3H6     7.3   1.2694     3296     4184

HTP/CH4O     3.2   1.1968     3102     3712
HTP/C2H6O    4.5   1.2458     3151     3926
HTP/C3H8O    5.2   1.2625     3167     3998
HTP/N2H4     2.2   1.2608     3283     4139
HTP/RP–1     7.3   1.3059     3223     4209
HTP/C6H6     6.6   1.3201     3210     4238
HTP/C3H4O    4.2   1.3010     3283     4271
HTP/C10H16   7.1   1.3190     3264     4305
HTP/C7H8     6.6   1.3496     3287     4436

F2/H2       14.6   0.6553     4704     3083
F2/H2O       2.1   1.2942     2876     3722
F2/HTP       0.88  1.4689     2966     4357
F2/NH3       3.4   1.1770     4115     4843
F2/B2H6      6.4   1.1314     4416     4996
F2/N2H4      2.3   1.3073     4212     5506

AZ50 (50% UDMH and 50% N2H4 by mass)
HTP (98% H2O2 and 2% H2O by mass)
Atsetam (32% C2H2 and 68% NH3 by mass)

        MP     BP
H2    -259.1 -252.9 Hydrogen
O2    -218.3 -182.9 Oxygen
CH4   -182.5 -161.5 Methane
B2H6  -164.9  -92.5 Diborane
C2H6  -182.8  -88.6 Ethane
C2H2   -82.2  -75   Acetelyne
C3H8  -187.6  -42.1 Propane
NH3    -77.3  -33.3 Ammonia
C3H6  -129    -33   Cyclopropane
C3H4  -102.7  -23.2 Methylacetylene
C4H6  -109     -4.5 1,3-Butadiene
C4H6  -119      2   Cyclobutene  HF?
N2O4   -15     21.2 Nitrogen Tetroxide
CH4O   -98     64.7 Methanol
C2H8N2 -57     54   Unsymmetrical Dimethylhydrazine (UDMH)
C2H6O -114     78   Ethanol
C6H6     5.5   80   Benzene
C3H8O  -89.5   82   Isopropanol
C7H8   -27     88   Quadricyclane
CH6N2  -52     91   Monomethylhydrazine (MMH)
H2O      0.0  100.0 Water
N2H4     1    113.5 Hydrazine
C3H4O  -53    114.5 Propargyl Alcohol
CH1.95 -62    147   RP-1
C10H16        158   Syntin


Aerozine50 density = 0.8818 kg/L

Efficiency = 97.4%
Chamber Pressure = 20.7 MPa
Expansion Ratio = 77.5
« Last Edit: 10/24/2014 08:20 am by Steven Pietrobon »
Akin's Laws of Spacecraft Design #1:  Engineering is done with numbers.  Analysis without numbers is only an opinion.

Offline Steven Pietrobon

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Re: RP-1, methane, impulse density
« Reply #42 on: 10/24/2014 08:20 am »
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.
Akin's Laws of Spacecraft Design #1:  Engineering is done with numbers.  Analysis without numbers is only an opinion.

Offline Proponent

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Re: RP-1, methane, impulse density
« Reply #43 on: 10/24/2014 09:01 am »
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, in some sense the combination is rewarded for being bulky.

EDIT: "I''m" -> "in" in final sentence.
« Last Edit: 10/25/2014 08:26 am by Proponent »

Offline R7

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Re: RP-1, methane, impulse density
« Reply #44 on: 10/24/2014 10:51 am »
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.

Smaller number -> smaller tanks -> smaller portion of the final mass is tank mass -> the better.
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Offline MP99

Re: RP-1, methane, impulse density
« Reply #45 on: 10/27/2014 08:18 am »


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.

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

Offline simonbp

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Re: RP-1, methane, impulse density
« Reply #46 on: 10/27/2014 12:13 pm »
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.

Offline Proponent

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Re: RP-1, methane, impulse density
« Reply #47 on: 10/27/2014 12:58 pm »
That's assuming constant wall thickness, but the mass of a pressure vessel scales as the product of pressure by volume for constant tensile strength.  This tends to suggest that volume normalised in some way is a valid metric.

As Whitehead points out (see page 8 of the attached paper), however, the mass of the tanks is by no means the whole story -- supporting structures can be quite significant too.

EDIT:  Added missing 'a' near end of second sentence.
« Last Edit: 10/28/2014 08:49 am by Proponent »

Offline baldusi

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Re: RP-1, methane, impulse density
« Reply #48 on: 10/27/2014 03:49 pm »


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.

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

Just 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\OrbitLEOGTO
LOX+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.

Offline Hyperion5

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Re: RP-1, methane, impulse density
« Reply #49 on: 10/27/2014 04:22 pm »
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.

Hypergolics are only the best if you don't factor in their terrible toxicity.  If you're looking for a non-toxic, high impulse density mix, it's hard to beat RP-1/H2O2 (95%+ Hydrogen Peroxide).  That mix is denser than almost any hypergolic mix, though its not-so-great Isp prevents it from winning out on impulse density.  On both cost of propellants and of complications, "keroxide", as Lobo has dubbed it, should win out.  Only problem with the mix for Spacex is that it freezes at water-like temperatures and you can't produce RP-1 on Mars (at least, not for a long while).  Which brings them right back to CH4. 



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.

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

Just 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\OrbitLEOGTO
LOX+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.
 

I just want to be clear on this.  Are you shrinking the propellant tanks or using the extra propellant within the same volume tanks to enable more powerful engines to fling the rockets skyward? 


Offline strangequark

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Re: RP-1, methane, impulse density
« Reply #50 on: 10/27/2014 08:39 pm »
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.

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.

Ah, so what you forgot is that the specific heat for fuel rich combustion products using methane is what matters, rather than methane itself. You set a combustion temperature, based on what your turbine can handle, then determine the appropriate O/F ratio to reach that temperature. The specific heat advantage is much more pronounced for the combustion products, both due to the T vs. O/F curve for methane, and the uniquely high hydrogen content. As a result, the trade between methane vs other LHCs is stronger for staged combustion cycles than something like expander.

Offline Proponent

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Re: RP-1, methane, impulse density
« Reply #51 on: 10/27/2014 10:30 pm »
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).

Offline R7

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Re: RP-1, methane, impulse density
« Reply #52 on: 10/30/2014 12:54 pm »
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).

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.

Attaching a capture what species RPA thinks a notional methane fuel-rich preburner contains (800K at injector, 100bar pressure, small 1.5 contraction, miniscule 1.0001 expansion because AIUI traditional turbomachinery does not operate with clearly supersonic flow). Substantial amount of free carbon and hydrogen.

Compare to similar but oxidizer rich case, very clean.
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Offline Proponent

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Re: RP-1, methane, impulse density
« Reply #53 on: 01/05/2015 12:04 pm »
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).

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.

Indeed, the RPA results you provide demonstrate that I was naive in assuming that the exhaust of a fuel-rich preburner would be essentially the fuel.  In my naive analysis, I pointed out that while methane's heat capacity of  3.99 kJ K-1 kg-1 at, e.g., 800 K is 13% higher than propane's (3.51 kJ K-1 kg-1), the greater volume of lox-methane propellant to be pumped results in a net advantage for propane.

So let's compare specific heats of fuel-rich preburner exhausts for lox-methane and lox-propane.  Repeating your calculation for lox-methane at 10 MPa and 800 K (attached), I too get a specific heat of 4.90 kJ K-1 kg-1.  A similar calculation (also attached) for lox-propane (O/F = 0.085) gives a specific heat of 4.40 kJ K-1 kg-1.  Now methane's advantage in specific heat is just 11% -- less than in the naive case.  In other words, considering more realistic preburner chemistry appears only to increase propane's advantage over methane from a thermodynamic viewpoint.

On the other hand, the output of the lox-propane preburner appears to be 30% solid carbon by mass.  Does anybody really want to run a turbine with that much solid carbon, which would then be fed into the main combustion chamber?  Sounds like a recipe for disaster to me.

So, that might seem to be a good reason to prefer methane to propane for staged combustion.  But ... have a look at the methane exhaust products: 8% solid carbon.  That's better than 30%, but it  still sounds bad.

The bottom line is that while I'm sure SpaceX and BO have good reasons for choosing methane for their staged-combustion engines, it's still not obvious to me that heat capacity is what makes methane superior to other light hydrocarbons.

EDIT: "I" -> "it" in penultimate sentence.
« Last Edit: 08/24/2015 06:01 am by Proponent »

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Re: RP-1, methane, impulse density
« Reply #54 on: 06/08/2015 08:57 am »
I found an interesting company called Synfuels International that is able to make Ethylene (C2H4), gasoline and jet fuel from alcohol. When I ran ethylene with the USAF Isp code, I got the following results.

Propellants  MR   dp (kg/L)  ve (m/s) Id (Ns/L)
O2/CH4       3.6   0.8376     3656     3062
O2/C2H4      2.7   0.9007     3678     3313
Efficiency = 97.4%
Chamber Pressure = 20.7 MPa
Expansion Ratio = 77.5

        MP     BP
O2    -218.3 -182.9 Oxygen
CH4   -182.5 -161.5 Methane
C2H4  -169.2 -103.7 Ethylene


Compare to methalox, ethylox (liquid ethylene and liquid oxygen) has a density that is 7.5% greater and an Isp that is 0.6% greater, giving an impulse density that is 8.2% greater!

On the way to make their jet fuel they make a butene (C4H8)/hexene (C6/H12) mixture. Unfortunately, I don't know what the density and heat of formation of this combination is, so I can't determine its performance.

Their jet fuel has a density of 0.803 kg/L. The paper below examined p-cymene (C10H14) and pinane (C10H18) which both have a density of 0.86 kg/L! That could be a great rocket fuel if it gives a good Isp. The freezing point of both chemicals is less than -70 C.

If anyone can help with heat of formations, carbon to hydrogen ratios and densities of these chemicals or mixtures, that would be much appreciated. Thanks.

Attached are some slides and a paper on their interesting brews.
« Last Edit: 06/08/2015 10:05 am by Steven Pietrobon »
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Re: RP-1, methane, impulse density
« Reply #55 on: 06/08/2015 10:32 am »
I found a heat of formation -174.7 kJ/mol for Pinane (C10H18) on the NIST website. I get the following results

Propellants  MR   dp (kg/L)  ve (m/s) Id (Ns/L)
O2/CH4       3.6   0.8376     3656     3062
O2/C2H4      2.7   0.9007     3678     3313
O2/RP–1      2.8   1.0307     3554     3663
O2/C10H16    2.6   1.0471     3608     3778
O2/C10H18    2.7   1.0531     3599     3790


C10H16 is Syntin, formerly used by Russia in their Soyuz vehicles. Compared to kerolox, Pinane/LOX has a 1.2% better Isp and 2.2% better density, giving a 3.5% better impulse density.
Akin's Laws of Spacecraft Design #1:  Engineering is done with numbers.  Analysis without numbers is only an opinion.

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Re: RP-1, methane, impulse density
« Reply #56 on: 06/08/2015 01:22 pm »
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.

Obviously I'm neither a chemist or a rocket engineer.
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Offline Impaler

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Re: RP-1, methane, impulse density
« Reply #57 on: 06/09/2015 03:22 am »
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.

Offline Steven Pietrobon

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Re: RP-1, methane, impulse density
« Reply #58 on: 06/09/2015 05:56 am »
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.

The differences in mass and volume flow rates would be such that is probably not practical. The engine has to be designed for the particular fuel used.
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Offline Burninate

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Re: RP-1, methane, impulse density
« Reply #59 on: 06/09/2015 09:14 am »
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.

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/
« Last Edit: 06/09/2015 09:23 am by Burninate »

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Re: RP-1, methane, impulse density
« Reply #60 on: 06/09/2015 09:27 am »
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? It's a major chemical feedstock due to capability to form chains with molecular weight up to millions to make plastics and whatnot. For this reason a colony will probably establish ethylene production anyway.

What's the route from basic Martian resources to synthesize ethylene?
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Offline Steven Pietrobon

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Re: RP-1, methane, impulse density
« Reply #61 on: 06/15/2015 06:00 am »
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?

You need a catalyst for ethylene to polymerise, so the liquid may still be stable under heat and pressure. According to

http://www.chemguide.co.uk/mechanisms/freerad/polym.html

you 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.

Quote
What's the route from basic Martian resources to synthesize ethylene?

Here is one possible way. First we have Sabatier

2CO2 + 8H2 → 2CH4 + 4H2O

Followed by water electrolysis

6H2O → 6H2 + 3O2

Followed by oxidative coupling of Methane

http://siluria.com/Technology/Oxidative_Coupling_of_Methane

2CH4  +  O2   →   C2H4 +  2H2O

The overall result is

2CO2 + 2H2 → C2H4 + 2O2

However, we need 2.37O2 for each C2H4, which means we need to generate an additional 0.37O2, which can be done using the reverse water reaction:

CO2 + H2 → CO + H2O

and then electrolysis

H2O → H2 + 0.5O2

That's a lot more efficient then directly electrolysing CO2 into CO and O2. My initial reaction is that the extra complexity is probably not worth it. However, the reduced amount of H2 from Earth plus the higher Isp and density of ethylox might possibly make it worthwhile.

This web site

http://www.marspedia.org/index.php?title=Reverse_Water-Gas_Shift_Reaction

says that you can use the CO from the reverse water reaction to also make methane and methanol, from which you can make ethylene.

There is also some research on CO2+H2 to C2H4, but I don't think it has been industrialised yet.
« Last Edit: 06/15/2015 06:38 am by Steven Pietrobon »
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Offline Proponent

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Re: RP-1, methane, impulse density
« Reply #62 on: 08/25/2015 06:03 am »
Methane may have an advantage in fuel-rich staged combustion, but I'm still looking for an aha moment to explain its current popularity in other cycles.

The attached plot shows the viscosities as functions of temperatures of methane, ethane and propane as solid curves in brown, red and orange, respectively (mneumonic: express the number of carbon atoms with the resistor color code).  Dotted red and orange lines show ethylene (ethene) and propylene (propene).  The solid line across the top shows dodecane, standing in for RP-1, at 20oC; RP-1 itself would probably be a little more viscous still.  All data are from NIST.

At lox temperatures, propane is molasses in January, if you'll pardon the hemispherism.  For the most part, though, it's less viscous than RP-1, which is often used in cooling.  So, is methane's very low viscosity, even at lox temperatures, a significant advantage, or is it just a marginal benefit?

Suppose you did want to use lox-propane with a common-bulkhead tank.  The propane near the bulkhead could be pretty sludgey, but how big a problem would that be?

Offline Proponent

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Re: RP-1, methane, impulse density
« Reply #63 on: 03/09/2016 02:15 pm »
When I previously posted in this thread, I was under the misapprehension that Blue Origin's BE-4 ran a fuel-rich staged-combustion cycle.  In fact, it's oxidizer-rich (fact sheet attached).  While reasons for SpaceX's Raptor to burn methane rather than a slightly heavier hydrocarbon might include:

  • 1. Relative cleanliness of combustion products on the fuel-rich side of Raptor's full-flow cycle;
  • 2. High mass-specific heat capacity, which appears to partially offset methane's low density;
  • 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
  • 4. Proximity of methane's boiling point to oxygen's, which allows common-bulkhead tanks;

  • only the last of these applies to the BE-4.  Even then, one could point out that propane is liquid at lox temperatures and typical ullage pressures.  However, it will be rather viscous.  Also, the use of a sub-cooled propellant makes autogenous pressurization trickier (do we know whether autogenous pressurization is planned for Vulcan, MCT or modified Falcon stages?).

    I suspect the choice of methane comes is an example of the fact that, in the paraphrased words of a forum member we're very lucky to have, "amateurs argue about engines; professionals argue about pressurization systems."

    Offline Robotbeat

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    Re: RP-1, methane, impulse density
    « Reply #64 on: 03/10/2016 02:52 am »
    Try subcooled propylene as a propellant. Even better than ethylene.
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    Offline Rei

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    Re: RP-1, methane, impulse density
    « Reply #65 on: 03/11/2016 09:39 pm »
  • 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

  • What about it doesn't make sense to you?  The methane fraction that comes out is much higher than that of other hydrocarbons.  You can customize it to some extent by controlling the process parameters - catalyst selection, heat, pressure, etc - but in general you get far more methane than other hydrocarbons.

    Now, if you want higher hydrocarbons and not the methane, you can always convert the methane (and other undesirable fractions) to syngas and then route it back in as a feedstock.  But it's added complexity, greater energy consumption, and lower throughput.

    Offline Proponent

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    Re: RP-1, methane, impulse density
    « Reply #66 on: 03/12/2016 07:43 am »
    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).

    Offline Robotbeat

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    Re: RP-1, methane, impulse density
    « Reply #67 on: 03/16/2016 07:25 am »
    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).
    Then you doubt incorrectly.
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    Offline Nilof

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    Re: RP-1, methane, impulse density
    « Reply #68 on: 03/16/2016 11:27 am »
    Another reason that Musk has mentioned at some point is that LNG/LOX is among the cheapest propellant combinations available. Sure, propellant costs are a tiny fraction of launch costs now, but in an architecture where you're aiming for 500 grand per seat to go to Mars the current market price of alternative propellants such as Ethylene suddenly becomes a rather significant portion of the total cost if you use it on every stage.

    So I think that Methane was chosen partly because it happens to be one of few propellants in the intersection of (high performance propellants that can easily be made on Mars) and (propellants with a reasonably low market price on Earth).
    For a variable Isp spacecraft running at constant power and constant acceleration, the mass ratio is linear in delta-v.   Δv = ve0(MR-1). Or equivalently: Δv = vef PMF. Also, this is energy-optimal for a fixed delta-v and mass ratio.

    Offline sevenperforce

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    Re: RP-1, methane, impulse density
    « Reply #69 on: 03/16/2016 03:26 pm »
    Try subcooled propylene as a propellant. Even better than ethylene.
    Wikipedia's propellant table lists ethylene as having a noticeably better specific impulse than methane. Of course, it has much higher density.

    It doesn't list propylene, though...and isn't propylene more likely to coke? I haven't been able to find anything more than 347 s for the vacuum isp of a propylene/LOX engine.

    For increased impulse density, what about using an H2O2/LOX blend during launch and then switching to LOX-only as you go higher up? Peroxide has the advantage of self-pumping, and it has a notably higher density than LOX with only a meager hit in specific impulse. It would also really improve T/W ratio at launch, reducing gravity drag.

    Offline Proponent

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    Re: RP-1, methane, impulse density
    « Reply #70 on: 03/16/2016 04:29 pm »
    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 Origin
    Liquefied natural gas enhances affordability and reusability

    Liquefied 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?

    Offline Rei

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    Re: RP-1, methane, impulse density
    « Reply #71 on: 03/16/2016 06:29 pm »
    You need a catalyst for ethylene to polymerise, so the liquid may still be stable under heat and pressure. According to

    http://www.chemguide.co.uk/mechanisms/freerad/polym.html

    you 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.

    There's a wide range of catalytic processes for ethylene; some operate at near atmospheric pressure.  The reaction rate is positively correlated with temperature.  I'm not sure how much of a problem it would be without a catalyst but at very high temperatures.  Probably not too mucy.  Polyethylene is considered to be a good fuel in its own right in some systems - for example, hybrid rockets.  It readily melts/vaporizes, like paraffin wax, so gets whipped into a spray in the flow which gives it a large reactive surface area (very important in hybrids)

    As for how it's produced: The production process to polyethylene first involves the sabatier reaction to produce methane, a second (or combined) step to produce simple alkenes, a filter stage to extract ethylene from the rest of the stream, an oxidation stage for the rest of the stream to turn it into syngas to feed back into the original process, a couple purification steps on the ethylene stream, and then a polymerization stage.  There's as mentioned a variety of processes involved in polymerization... widely varied pressures, but generally fairly low temperatures required.  The problem with going too hot is that it catalyzes too fast and gunks up your reactor; since polymerization is exothermic, you can even (particularly in gas-phase polymerization systems) get runaway polymerization that basically turns your catalyst bed into a big block of plastic  ;)   Most processes (although not gas phase) end up with the catalyst embedded in the output product, which may or may not be extracted, and to varying degrees, and with varying levels of consumables involved.  Even in gas-phase reactors, you still need to renew/replace your catalyst over time.

    Rather than just filtering out only ethylene from the sabatier output stream, though, better would be to distill multiple products out, because each is useful.  For example, propylene reacts with ammonia (Haber process) and oxygen in the SOHIO process to produce a stream that you wash with sulfuric acid to produce a bit of N2 waste gas, some NH3SO4, some HCN, some CH3CN, but mainly acrylonitrile.  Polymerizing *that* yields PAN, which can be gel spun to strong PAN fiber.  This in turn can be first oxidized, then carbonized, followed up by a sulfuric acid surface treatment, to produce carbon fiber.   Of course, fibers alone, regardless of their composition, are only part of the story, you then generally want to produce them into weaves, and you still need resins to make composite parts... and manufacture would be quite difficult if you had to work in a space suit...

    Offline Proponent

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    Re: RP-1, methane, impulse density
    « Reply #72 on: 03/17/2016 06:56 am »
    I presume, then, the suitability of propylene, ethylene or other unsaturated hydrocarbons depends on the details.  They're probably not such good coolants, and it sounds like fuel-rich or full-flow staged combustion might be a disaster.  On the other hand, if cooling and pumping requirements are not placed on them, then they could be pretty attractive.  The one actual use of propylene I'm aware of is in some of Garvey Spacecraft Corporation's rockets (hopefully leading to nanosat launchers), which I believe are pressure fed -- don't know about the cooling.  Do we know of any actual uses of ethylene (or other uses of propylene)?

    Offline Rei

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    Re: RP-1, methane, impulse density
    « Reply #73 on: 03/18/2016 08:02 am »
    It depends on what your primary needs in a coolant are.  Like most rocket propellants, they have high specific heats.  Yes, my main concerns would be anywhere that they're being heated in small channels where there's no oxidizer present.  That said, there's all sorts of things one can do to prevent polymerization.  For example, the process is impurity-sensitive; in the event of runaway polymerization the solution is often simply to "poison" the reaction by injecting carbon monoxide or other hydrocarbons.  I'm sure a reasonable approach could be developed.

    I'm not aware of any full-scale liquid fuel systems that burn ethylene or propylene, and haven't even read about experimental scale ones, although there might have been some.  I have however read about experimental scale usage of polyethylene as a hybrid rocket fuel.  Aluminized paraffin wax and aluminized polyethylene appear to be the best hybrid rocket fuels out there: cheap, stable, easy to cast, easy to melt/vaporize, burn readily with LOX, and low viscosity upon heating, so easy to whip into a fast-burning mist.  You get significantly higher ISP and propellant density with aluminized paraffin or polyethylene burned with LOX than you do with LOX/RP-1 (mainly due to the aluminum).  A lot of hybrid work had previously focused on polybutadiene as the fuel, which is very mature in its use as a solid fuel binder - but hybrids are not solids, and polybutadiene isn't as readily mobilized and thus gets a significantly lower burn rate, thus requiring significantly more combustion channels, thus leading to less burn stability.

    Offline Rei

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    Re: RP-1, methane, impulse density
    « Reply #74 on: 03/18/2016 08:08 am »
    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).
    « Last Edit: 03/18/2016 08:09 am by Rei »

    Offline notsorandom

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    Re: RP-1, methane, impulse density
    « Reply #75 on: 03/18/2016 12:12 pm »
    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 Origin
    Liquefied natural gas enhances affordability and reusability

    Liquefied 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?
    Methane seems to have a lot of attractive qualities and its drawbacks are managed easily. Density and IPS are important but there are other factors as well that are important. Things like pressurization, thermal capacity, viscosity, and coking. It could be that Methane is the ultimate compromise. Other propellant combinations offer better performance in some of those categories but worse in others. That it is easy to make on Mars may just be the cherry on top.

    Offline hkultala

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    Re: RP-1, methane, impulse density
    « Reply #76 on: 04/14/2016 09:01 pm »
    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=4.8), while impulse density peaks at  633 kg/m3 (O/F=17.8).  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. But increasing the amount of hydrogen by factor of 3 is the normal way of doing hydrolox engines, only less than 1/3 of all the hydrogen is burning.

    But optimizing for isp by using a lot of extra hydrogen also makes the impulse density very bad as the extra hydrogen still consumes a lot of space.

    Most air-lit/upper stage engines(RL-10,J-2) use mixture ratio 5.5 and ground-lit/sustainer or booster engines use mixture ratio of about 6(RS-25, RS-68).

    It would seem that RS-68 uses too small mixture ratio;

    Raising the mixture ratio from 6 to 7 might decrease isp by about 2%, but it would increase impulse density (and thrust) of same-sized engine by 10%. However, for a 5.5km/s rocket stage and ~400s engine, the 2% isp drop means that only about 3% more propellant is needed. So Delta IV could manage with ~7 % smaller tanks and lighter engines if the mixture ratio would b 7 instead of 6.

    Raising it to 8 instead would decrease isp about 4%, and impulse density and engine T/W would increase by about 17%. 4% of lost isp would cost only about 6% fuel for the 5.5 km/s stage, so about 10% smaller tanks and lighter engines would do.

    So what am I missing here? Why did they make RS-68 to use se low mixture ratio?


    Going much more 8 does not seem to make much sense as after 8 isp start to decrease by a lot and the impulse density increase starts to get much smaller after about 7.5.




    « Last Edit: 04/14/2016 09:01 pm by hkultala »

    Offline TrevorMonty

    Re: RP-1, methane, impulse density
    « Reply #77 on: 04/14/2016 09:31 pm »
    This from Jeff Bezo interview. I imagine SpaceX are thinking along same lines hence Raptor. NB BE4 uses LNG (typically 95% methane) so it is cheaper than pure methane.

    http://www.geekwire.com/2016/interview-jeff-bezos/

     It’s why we are using liquid natural gas [for the BE-4 engine], because our goal is to make spaceflight so cheap that the cost of the fuel actually matters.

    Right now, the cost of the propellant is minimal compared to throwing the hardware away. But once we can get to a place where we are not throwing the hardware away, and we have real reusability, then we want to be using a fuel that is very low-cost. And nothing is lower-cost than liquid natural gas.



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    Offline Proponent

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    Re: RP-1, methane, impulse density
    « Reply #78 on: 04/14/2016 10:12 pm »
    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)0.23

    for 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.
    « Last Edit: 04/15/2016 08:01 am by Proponent »

    Offline Proponent

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    Re: RP-1, methane, impulse density
    « Reply #79 on: 04/14/2016 10:20 pm »
    Good point. For perfect burning, mixture ratio should be 16.

    You mean 8, right?

    Quote
    It would seem that RS-68 uses too small mixture ratio;

    Maybe it's partly to increase the thrust-to-weight ratio.

    Offline hkultala

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    Re: RP-1, methane, impulse density
    « Reply #80 on: 04/15/2016 10:21 am »
    Good point. For perfect burning, mixture ratio should be 16.

    You mean 8, right?

    Whoops, missed some multiplication by 2 in my chemistry.
    Quote
    Quote
    It would seem that RS-68 uses too small mixture ratio;

    Maybe it's partly to increase the thrust-to-weight ratio.

    Makes T/W ratio of the rocket worse without solids, as the thrust should be about linearily propotional to the propellant density, and it increases more than the propellant weight increases.

    2 possibilities come into mind:

    1) It's optimized to be used with many solid boosters, not without solids, and with solids the solids have to lift less first stage propellant mass with lower mixture ratios
    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?
    « Last Edit: 04/15/2016 10:25 am by hkultala »

    Offline RanulfC

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    Re: RP-1, methane, impulse density
    « Reply #81 on: 04/15/2016 01:31 pm »
    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).

    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.

    Did I miss a cite somewhere?

    Randy
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    Offline RanulfC

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    Re: RP-1, methane, impulse density
    « Reply #82 on: 04/15/2016 02:06 pm »
    To further develop my understanding of the trade between specific impulse and density, I've done a little thought experiment on ground-launch stages.
    <snipping some excellent stuff>

    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?

    Quote
    1.  Except for JP-5 (the composition of which I don't know),

    JP-5 or JP-10? JP-5 Material Safety Data Sheets, (MSDS) are available on-line and composition properties sheets IIRC:
    http://www.henrycoema.org/LEPC/LEPCPlan/Apdx%20E-Other%20Information/JP5_9942_clr.pdf
    http://www.cpchem.com/msds/100000014588_SDS_US_EN.PDF
    http://www.atsdr.cdc.gov/toxprofiles/tp121-c3.pdf

    JP10's a bit more difficult to find:
    http://www.boulder.nist.gov/div838/SelectedPubs/IR%206640%20ms.pdf
    http://www.madsci.org/posts/archives/2000-08/966261622.Ch.r.html

    If that helps.

    Randy
    From The Amazing Catstronaut on the Black Arrow LV:
    British physics, old chap. It's undignified to belch flames and effluvia all over the pad, what. A true gentlemen's orbital conveyance lifts itself into the air unostentatiously, with the minimum of spectacle and a modicum of grace. Not like our American cousins' launch vehicles, eh?

    Offline R7

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    Re: RP-1, methane, impulse density
    « Reply #83 on: 04/15/2016 02:58 pm »
    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?

    Going from O/F 6 to 7

    - increases Tc about 100K
    - reduces the available coolant franction of total propellant from 1/7 to 1/8 (12.5% less)
    - increase the free oxygen species in combustion gases by an order of magnitude.

    Even alone none of the three are good news for engine developer. All three together compound each others' problems. Higher Tc would require more cooling and increased oxygen content may require protective oxide layer ... which reduces thermal conductivity ... which requires even more aggressive cooling ... using reduced amount of coolant.

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    Offline Proponent

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    Re: RP-1, methane, impulse density
    « Reply #84 on: 04/15/2016 05:36 pm »
    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.

    Attached are liquid viscosities as a function of temperature, from NIST.  I left hydrogen off the plot to avoid compressing the hydrocarbon curves.  Take my word for it, hydrogen's viscosity is pretty low.  The dotted horizontal lines show RP-1's viscosity at a typical temperature of 21oC and at 20oF (266 K), which I believe is the temperature to which SpaceX chilled its RP-1 on CRS-7.  I could imagine that sub-cooling propane too much could cause a viscosity problem (if you're not using viscosity to your advantage, as Rei suggests).  On the other hand, maybe viscosity of propane in the tank doesn't matter too much and it warms up anyway before reaching any turbomachinery or injector, where high viscosity really would be a problem.

    I don't have any data for the more exotic fuels.  Cyclopropane is probably quite inviscid even at low temperatures, and it should have a good specific impulse (probably something approaching syntin's, which is three cyclopropane rings and a methyl group stuck together), though it is expensive.

    EDIT:  Changed musing on possible unimportance of low-temperature viscosity.  Corrected temperature of chilled RP-1.
    « Last Edit: 04/16/2016 03:26 am by Proponent »

    Offline Proponent

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    Re: RP-1, methane, impulse density
    « Reply #85 on: 04/15/2016 05:47 pm »
    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?

    Yes, absolutely (though whether the operations costs are just slightly higher or not might be another question).  In fact ideally the mixture ratio of each stage changes during the burn.  At lift-off, the ideal mixture ratio in principle is that corresponding to maximum impulse density.

    Quote
    JP-5 or JP-10? JP-5 Material Safety Data Sheets, (MSDS) are available on-line and composition properties sheets IIRC ....

    Thanks.  What I learn is that JP-5 too is basically kerosene with lots of additives.  I should have guessed that. So, following my scheme, it would have the same color (red) as RP-1.
    « Last Edit: 04/15/2016 05:51 pm by Proponent »

    Offline Hyperion5

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    Re: RP-1, methane, impulse density
    « Reply #86 on: 04/16/2016 01:10 am »
    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.

    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?

    Offline Proponent

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    Re: RP-1, methane, impulse density
    « Reply #87 on: 04/16/2016 01:48 pm »
    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?

    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.

    Offline S.Paulissen

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    Re: RP-1, methane, impulse density
    « Reply #88 on: 04/17/2016 05:36 pm »
    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?

    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.

    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. 

    I only ask because I've been designing a very amateur-style non-flight pressure-fed rocket motor with kerosene/N2O and found through my amateurish calculations expect a high 240s-range sea-level ISP with autopressurization of the N2O and CO2 pressurization of the kerosene. With helium etc, I would expect quite a bit more; enough to be useful in ISRU situations that don't have TEA-TEB at the ready.
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    Offline Proponent

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    Re: RP-1, methane, impulse density
    « Reply #89 on: 04/18/2016 12:56 pm »
    ... the thrust should be about linearily propotional to the propellant density ....

    Why density?  I would think impulse density most relevant.

    Offline Rei

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    Re: RP-1, methane, impulse density
    « Reply #90 on: 04/19/2016 11:47 pm »
    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. 

    It really is.  A typical N2O monoprop will get an Isp around 170 or so.  That's just really, really bad.  Fine for an RCS or small craft's OMS, maybe primary propulsion for a small probe... but you definitely don't want that for a launch vehicle.

    Quote
    through my amateurish calculations expect a high 240s-range sea-level ISP

    One, 240 is still a very bad ISP for a launch vehicle, even sea level.  And two, no, you're not going to get that sort of ISP.  Oh great, now I need to dig out CEA2... let's see.. I'll be generous and give it 200 bar chamber pressure, let's keep the pi/pe figures and mdot the same as from the SSME, basically giving you an SSME-ish N2O monoprop engine... okay, here you go:

    Quote
    *******************************************************************************

             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

    That's a vacuum ISP of 203.  From one heck of an advanced N2O engine!  ;)

    (CEA2's ISP figures are a bit wierd... one, they're not divided by 9,81, and two the one labeled "Ivac" assumes an infinite length nozzle; in this setup, the Isp is also vacuum).  Also note that CEA2 is generous to begin with; it's calculating equilibrium exhaust properties, there's no frozen combustion involved.

    As for returning from LMO to surface in a repeatable manner (the post before yours)... Mars' atmosphere is thin but it's not *that* thin, you can still aerobrake in it just fine.  The main disadvantage vs. Earth is the need for a powered landing.  Also parachutes would be trickier if you want them to be reusable - not only do they have to survive just fine, but be repacked by people in space suits, after having gotten covered in Martian dust and potentially landing on sharp rocks.  You can limit your aerobraking to an aeroshell without any parachutes, but then that means even more propellant needed for landing.

    I once ran the numbers up the delta-V requirements for powered landing on Mars, with and without parachutes, but I can't be bothered to dig it up right now.  ;)
    « Last Edit: 04/19/2016 11:59 pm by Rei »

    Offline S.Paulissen

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    Re: RP-1, methane, impulse density
    « Reply #91 on: 04/20/2016 02:09 am »
    Hah, thanks for that,  and most certainly is ghastly.  But I was not intending to use it ONLY as a mono-prop, I did not do a good enough job to get that across, my bad.  And you are certainly right with 240, being bad, but my number was what seemed achievable with a pressure fed system bi-propellant built by an amateur on a very very low budget, not something you'd launch into space.

    I was thinking more along the lines of using the N2O as a monoprop for RCS on a vehicle also using it as an oxidizer with a real fuel like methane or kerosene (likely to former rather than the latter as it's an ISRU fuel).

    EDIT: Thank you for pointing me at CEA2.  I cracked it open and put the numbers in as if it was the bi-propellant system I was intending and got 314s at sea level and 325s in vacuum.  I guess that answers my question.
    « Last Edit: 04/20/2016 03:07 am by S.Paulissen »
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    Offline Robotbeat

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    Re: RP-1, methane, impulse density
    « Reply #92 on: 04/26/2016 07:49 pm »
    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)0.23

    for 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.
    Bravo! Impressive analysis. Now can you try subchilling the propellants like Dunn?
    Chris  Whoever loves correction loves knowledge, but he who hates reproof is stupid.

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    Offline Rei

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    Re: RP-1, methane, impulse density
    « Reply #93 on: 04/26/2016 10:46 pm »
    Quote
    , tanks weigh 10 kg/m3

    Can 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.

    Quote
    pressurants and residuals are each 0.25% of the initial propellant load

    Again, is this defensible?  I certainly wouldn't expect highly voltatile cryogenics to have the same residual as far more viscous, non-volatile liquids.

    Quote
    Chamber pressure is 20 MPa and the area expansion ratio is 40:1.

    Assuming a perfect burn, no frozen combustion?

    Quote
    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.

    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)

    Quote
    But, 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)0.23

    I 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.

    That's not to say that there's no room for improvement - there absolutely is.  But the basics here are no mystery.  Unless you're using solids or you want to destroy the environment and human health by using a fluorine-containing oxidizer, the only obvious choice is LOX - it just performs so much better than its competitors, and its properties are fairly tame (by oxidizer standards, at least... which isn't saying much).  And hydrogen is in a league of its own for upper stages - unless you want to burn it with lithium in a triprop (or totally implausible combinations involving beryllium or boron), there's not much room for improvement.    With the exception of solids, the only real question is "what hydrocarbon do you want to burn with LOX on the first stage, and do you want to work aluminum into the mix?"

    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.tif

    You simply can't get a plausible structural coefficient with a low ISP propellant mix.  It just doesn't work.
    « Last Edit: 04/26/2016 10:51 pm by Rei »

    Offline Robotbeat

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    Re: RP-1, methane, impulse density
    « Reply #94 on: 04/27/2016 12:22 am »
    Quote
    , tanks weigh 10 kg/m3

    Can 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....
    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, it's appropriate to consider tank masses using the pressure vessel equation. And the pressure vessel equation indeed gives you tank masses with units of kg/m^3, regardless of scale.

    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.

    There are two fairly easy ways to reduce tank mass: Use materials with higher strength-to-weight ratio or operate at lower ullage pressure. This last one is perhaps the biggest reason why pump-fed rocket engines are used instead of pressure-fed.
    « Last Edit: 04/27/2016 12:23 am by Robotbeat »
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    Offline Robotbeat

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    Re: RP-1, methane, impulse density
    « Reply #95 on: 04/27/2016 12:35 am »
    ...

    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.tif

    You simply can't get a plausible structural coefficient with a low ISP propellant mix.  It just doesn't work.
    STRONGLY disagree. SSTOs are more dry-mass-dependent, and hydrogen's big dry mass penalty is significantly reduced by staging, so if the dry-mass situation looks bad with TSTO, it looks far worse with SSTO.

    I think dry mass and total rocket volume is a better stand-in for cost than lift-off mass is, and the only reason you'd use hydrogen for SSTO is if you're trying to minimize lift-off mass. Otherwise, you're far better with another propellant combination.
    Chris  Whoever loves correction loves knowledge, but he who hates reproof is stupid.

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    Offline Robotbeat

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    Re: RP-1, methane, impulse density
    « Reply #96 on: 04/27/2016 12:37 am »
    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.htm

    Hydrogen 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.
    « Last Edit: 04/27/2016 12:46 am by Robotbeat »
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    Offline Rei

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    Re: RP-1, methane, impulse density
    « Reply #97 on: 04/27/2016 01:53 am »
    Quote
    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

    No, 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.

    Quote
    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.

    And as was just demonstrated above, the Saturn-V first stage is half of an order of magnitude heavier per unit volume, not lighter. 

    Quote
    This 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".

    Quote
    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.htm

    Hydrogen 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:

    Quote
    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. 
    « Last Edit: 04/27/2016 02:00 am by Rei »

    Offline Robotbeat

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    Re: RP-1, methane, impulse density
    « Reply #98 on: 04/27/2016 01:59 am »
    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.

    You have to make some simplifying assumptions, and that's a pretty good one to make.
    « Last Edit: 04/27/2016 02:04 am by Robotbeat »
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    Offline Rei

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    Re: RP-1, methane, impulse density
    « Reply #99 on: 04/27/2016 02:17 am »
    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."

    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_Atlas

    Falcon 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.  Say, at 250 tonnes O2.  3,66m diameter = 10,5m cross section.  Thus stress on the bottom of the tank from G forces at launch is around 467kPa just from the fuel.   Now, you could call withstanding 467kPa a "pressure vessel", but it's no more a pressure vessel than any large cylindrical tank.

    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.

    Quote
    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.

    Okay, okay, let's see, Russia's probably the place to look to for modern, non-hydrogen upper stages... let's say, Proton-M's two uppermost stages?   That's a much more modern rocket than Saturn V, so you can't complain that the Saturn V would be somehow technologically more advanced.  I don't have exact tank sizes but I have propellant masses, and the fuel combination is generally 1,18g/cc.  Stage 3: 4185kg, ~34 m^3; tankage  123kg/m^3.  Briz-M: 2370kg; tankage volume, ballpark 17 m^3; ratio: ~139kg/m^3

    It just makes the case for some sort of "constant ratio" even worse.  Want me to make it even worse and compare to an actual modern hydrogen upper stage?

    It's just not even remotely realistic to act like there's some sort of fixed ratio.

    Now, you can make some sort of general function m(p,v,s) where:

     m = dry mass
     p = propellant scalar factor
     v = propellant volume
     s = stage factor (stages that start their burn further (in terms of delta-V) into the flight have a different ratio of engine mass to tank than lower stages)

    Where:
    m(p, v, s) = a * p * v^b * s^c

    ... where a, b,  and c are solved for by regression fitting and p for each propellant combination is determined empirically.  But of course you'll find that newer rockets tend to go under that curve and older rockets over it, so it might be worth adding in a "technology level scalar" in there as well.


    « Last Edit: 04/27/2016 01:48 pm by Rei »

    Offline dkovacic

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    Re: RP-1, methane, impulse density
    « Reply #100 on: 04/27/2016 03:18 am »
    Quote
    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

    No, 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.

    Quote
    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.

    And as was just demonstrated above, the Saturn-V first stage is half of an order of magnitude heavier per unit volume, not lighter. 

    Quote
    This 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".

    Quote
    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.htm

    Hydrogen 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:

    Quote
    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. 
    @Rei, I don't know from where you got your numbers for Saturn V. From Wikipedia, https://en.wikipedia.org/wiki/S-IC, first stage had roughly 2075m3 of propellant volume (you took only the oxidizer value), and empty mass of the second stage was 40 tonnes (see https://en.wikipedia.org/wiki/Saturn_V. Next, here is the quote from the Wikipedia article:

    Quote
    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.

    Also take into consideration that the 1st stage had to provide eight times more thrust and support roughly three times more mass. Also much more attention was given to optimization of second and third stage.

    So you should not compare first and second stage. Proper comparison would be to design hydrolox 1st stage of equivalent total impulse capability (lets assume you could achieve needed thrust with hydrolox). Even at 30kg/m3, you would need a lot more volume than 2000m3. Higher ISP gets you roughly 20% more impulse. But you would need almost 2.5x more volume. Add to that additional support mass for upper stages and weight of additional J-2 engines needed to give you adequate thrust.

    The beauty and insight of the Dr. Dunn's paper is that it gets you to view rocket performance from constrained volume perspective, not from mass perspective. If you rewrite rocket equation to express your payload mass as a function of Isp and propellant density (you can see detailed analysis at https://lilibots.blogspot.com/2016/02/rocket-equation-revisited.html), you can see that density has a huge impact on rocket performance. Thus hydrolox is not only worse then kerolox, it is even worse than HTPB (in case of pressure-fed design).

    So the proper question is why NASA kept on developing hydrolox technology. Origins go the sixties, and I am sure these guys were very smart. But they did not have Internet, NSF and Google back then. Notice that Dr. Dunn's paper, which would probably be long forgotten in the era before Internet, just keeps being quoted on and on in various places. Its conclusions are not relevant just for SSTO, they are relevant for any stage.

    Another view can be comparison of Saturn V and N1 rockets. On the first glance, they have roughly the same lift-off weight but N1 had half of TLI capacity. So you might conclude that Saturn V was simply more efficient due to hydrolox on upper stages. But it you take total propellant volume as a measure, then N1 appears close or even better than Saturn V.
    « Last Edit: 04/27/2016 04:15 am by dkovacic »

    Offline Robotbeat

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    Re: RP-1, methane, impulse density
    « Reply #101 on: 04/27/2016 03:20 am »
    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."

    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_Atlas

    Falcon 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. ...
    Way too high. T/W of the whole rocket is more like 1.15-1.4. And interestingly, the pressure at the bottom of the tank barely changes, basically staying the same as it is at launch (or lower, as you throttle-down near max-q and near the end of the stage burn). If you consider zero payload and massless tanks and a cylindrical tank (or constant cross-section), if you keep the rocket engine thrust constant, the pressure at the bottom will actually be constant, with the increased acceleration due to the emptying of the tanks being exactly countered by the reduction in propellant mass.
    Chris  Whoever loves correction loves knowledge, but he who hates reproof is stupid.

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    Offline Proponent

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    Re: RP-1, methane, impulse density
    « Reply #102 on: 04/27/2016 08:42 am »
    Now can you try subchilling the propellants like Dunn?

    That's definitely on my list of things to do, along with looking at upper stages.  There are a couple of things I'm pondering about treating sub-cooling.  Obviously, sub-cooling allows you to pack more propellant into tanks of a given size, but does it also improve the engine's T/W?  I'm thinking the best simple approximation is that it does not, because the propellants will have warmed up quite a bit by the time they've got very far into the engine (which includes all of the turbo machinery).

    The other thing is, what's a reasonable limit to impose on a propellant's viscosity?  Viscosity is probably most important at the injector, but by the time propellant gets there it will have warmed up quite a bit.  We know SpaceX chills RP-1 to 20 oF, where it's viscosity is about 3.3 cP.  Is it OK to chill propane all the way to a viscosity of 10 cP?  RP-1 to 20 cP?  JP-10 to 200 cP?

    Offline Proponent

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    Re: RP-1, methane, impulse density
    « Reply #103 on: 04/27/2016 10:01 am »
    ...

    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.

    Personally I think the flimsiest part of my analysis is my claim that what I define as the "available mass" is largely propellant-independent.  As Whitehead points out, "non-tankage structures are historically heavy," and to the extent that they support flight loads may depend on propellant mass more than I have allowed for.

    Quote
    Assuming a perfect burn, no frozen combustion?

    The specific-impulse figures assume shifting equilibrium.
     
    Quote
    Quote
    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.

    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)

    The criterion is the available mass fraction, which is shown by the grey contour lines in the figures.  Note that it's those contours, not the density and I[/i]sp, that change from plot to plot.

    Quote
    Quote
    But, 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)

    Good point, though the relatively low delta-V requirement for a martian SSTO means that hydrogen's not really at its best anyway.

    Quote
    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.

    Look at the tables.

    Quote
    Quote
        (specific impulse)(bulk density)0.23

    I can see no justification for the usage of a formula involving a linear multiplication of specific impulse and bulk density.

    It's simply an approximate figure of merit.  Both specific impulse and density matter, but density matters less; the exponent quantifies how much less.  To put it another way, if we were to show contours of constant available mass fraction on a log-log plot of density and specific impulse, the slope would be about 0.23 in the region of the plot occupied by the better propellant combinations.  A 1% boost in density is about as valuable as a 0.23% boost in specific impulse.

    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.

    Quote
    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.

    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.

    Offline Rei

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    Re: RP-1, methane, impulse density
    « Reply #104 on: 04/27/2016 11:56 am »
    ...

    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.

    It doesn't really matter what justifications he presents, in the real world, it demonstrably does not work like that.  You cannot simply multiply the volume times a scalar to get the tankage mass.

    Skimming over his (apparently also non-peer-reviewed) paper, it relies heavily on the Shuttle for most of its justifications, which is anything but an ordinary rocket design, and as a result anything but mass efficient (indeed, I didn't even consider it worthy of consideration above).  He apparently comes up with figures of LOX/LH stages being three times heavier per unit fuel mass.  That is demonstrably not in line with real figures (see the post above).  A typical LOX/LH stage (not just tank, but whole stage) has about 9 times the propellant mass as its dry mass.  A typical LOX/RP1 stage has about 10-15.  SpaceX is exceptional with 18,4 on the Falcon 9 first stage.  But even that isn't close to a threefold difference.  And many of the things that help make the Falcon 9 tanks light can also be applied to LOX/LH.

    Quote
    Quote
    Assuming a perfect burn, no frozen combustion?

    The specific-impulse figures assume shifting equilibrium.

    Okay, because real-world rocket engines don't operate at equilibrium.
     
    Quote
    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.

    Not at all - I fully respect the importance of density.  I just strongly disagree - based on the evidence from real world systems - that the density advantage of hydrocarbons overcomes the ISP advantage of hydrogen.

    That said, I think aluminized hydrocarbons might potentially give LOX/LH a run for their money.  That boosts the density even higher and raises the ISP significantly.  It also gives for a more stable burn.  The increased erosion rate might be a problem in a SSTO context, however.

    Quote
    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.

    Fair enough  :)

    Quote
    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.

    Of course, these sorts of factors go both ways.  In the context of a SSTO, a large tankage size increases the reentry cross section and thus lowers reentry heat loading per square meter. It also makes it easier to get more lift out of a design as needed.

    I'm not sure what you mean by "its high specific impulse reduces acceleration". Its low density reduces acceleration for the same turbopump throughput, not the high specific impulse.  Indeed, one of the disadvantages for LOX/LH for lower stages is that it takes significantly larger engines to produce the same amount of thrust as LOX/RP1.  This isn't as applicable for upper stages, where ISP becomes the much more important factor.
    « Last Edit: 04/27/2016 12:02 pm by Rei »

    Offline Rei

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    Re: RP-1, methane, impulse density
    « Reply #105 on: 04/27/2016 01:22 pm »
    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/m³

    LOX/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.
    « Last Edit: 04/27/2016 01:36 pm by Rei »

    Offline dkovacic

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    Re: RP-1, methane, impulse density
    « Reply #106 on: 04/27/2016 02:54 pm »
    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/m³

    LOX/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.
    Well, I checked Ariane V ECA stage. According to wikipedia, it has dry mass of 14,700 kg. Vulcain 2 is 1.300kg of that mass (less than 9%), giving roughly 0.12g contribution at liftoff. Total propellant volume is 510m3 according to http://spaceflight101.com/spacerockets/ariane-5-eca/.

    On the other hand, Atlas V 1st stage has dry mass of 21,054 kg. RD-180 is 5,480 kg (around 26%), giving roughly 1.16g contribution at liftoff. Volume is 275m3 giving it ratio of 76kg/m3. So scaling to the same T/W would be close to 60kg/m3.

    Another significant difference is that Ariane V uses common bulkhead design, while Atlas V does not on CCB. According to Saturn V article on Wikipedia, that has major effect on mass, but it is hard to estimate how much really. It is generally not used for RP-1/LOX tanks. Falcon 9 seems to be exception here, which might explain why their kg/m3 metrics appear to be better.

    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.

    Offline Proponent

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    Re: RP-1, methane, impulse density
    « Reply #107 on: 04/27/2016 03:18 pm »
    All serious efforts toward SSTOs have used hydrogen.

    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 JP-4.

    Quote
    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:

    Quote
    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.

    A good reason for the constant-volume constraint: vehicles of similar size might be expected to have similar costs.  Hence, if one's goal is cheap SSTO rather than mass-efficient SSTO, assuming constant volume is not outrageous as a first approximation.  But I do agree that Dunn's treatment of hydrogen is rather ad hoc.  Once again, I recommend Whitehead's paper, which treats hydrogen more consistently but reaches a conclusion similar to Dunn's.

    Following is a more detailed comparison of winged lox-hydrogen and lox-kerosene SSTOs done by Burnside-Clapp in 1997, following.  It too winds up favoring the hydrocarbon.



    sci.space.policy > A LO2/kerosene SSTO rocket design, w/o AOL
       
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    A LO2/Kerosene SSTO Rocket Design (long)

    Mitchell Burnside Clapp
    Pioneer Rocketplane

    (view with a fixed pitch font such as courier or monaco)

    Abstract

    The NASA Access to Space LO2/hydrogen single stage to orbit rocket was examined, and the configuration reaccomplished with LO2/kerosene as the propellants. Four major changes were made in assumptions. First, the aerodynamic configuration was changed from a wing with winglets to a swept wing with vertical tail. The delta-V for ascent was as a result recalculated, yielding a lower value due to different values for drag and gravity losses. The engines were changed to LO2/kerosene burning NK-33 engines, which have a much lower Isp than SSME-type engines used in the access to space study, but also have a much higher thrust-to-weight ratio. The orbital maneuvering system on the Access to Space Vehicle was replaced with a pump-fed system based on the D-58 engine used for that purpose now on Proton stage 4 and Buran. Finally, the wing of the vehicle was allowed to be wet with fuel, which is a reasonable practice with kerosene but more controversial with oxygen or hydrogen. Additionally, in order to reduce the technology development needed, the unit weights of the tankage were allowed to increase by 17 percent.

    After the design was closed and all the weights recalculated, the empty weight of the LO2/kerosene vehicle was 35.6% lighter than its hydrogen fuelled counterpart.

    Introduction

    NASA completed a study in 1993 called Access to Space, the purpose of which was to consider what sort of vehicle should be operated to meet civil space needs in the future. The study had three teams to evaluate three different broad categories of options. The Option 3 team eventually settled on a configuration called the SSTO/R. This vehicle was a LO2/hydrogen vertical takeoff horizontal landing rocket. The mission of the Access to Space vehicle was to place a 25,000 pound payload in a 220 n.mi. orbit inclined at 51.6 degrees. The vehicle had a gross liftoff weight of about 2.35 million pounds. The thrust at liftoff was 2.95 million pounds, for a takeoff thrust to weight ratio of 1.2. The empty weight of the vehicle was 222,582 pounds, and the propellant mass fraction (defined here as [GLOW-empty]/GLOW) was 90.5%.

    Main power for this vehicle was provided by seven SSME derivative engines, with the nozzle expansion ratio reduced to 50. This resulted in an Isp reduction from 454 to 447.3 seconds. Each engine weighed 6,790 lbs, for an engine sea level thrust to weight ratio of 62.

    Aerodynamically the vehicle was fairly squat, with a fineness ratio (length:diameter) of 5. The overall length of the vehicle was 173 feet and its diameter was 34.6 feet. It had a single main wing (dry of all propellants) of about 4,200 square feet total area, augmented by winglets for directional control at reentry. The landing wing loading was about 60 lb/ft2. The oxygen tank was in the nose section. The payload was mounted transversely between the oxygen and hydrogen tanks, and was 15 feet in diameter and 30 feet long.

    This design exercise was among the most thorough ever conducted of a single stage to orbit LO2/LH2 VTHL rocket. It was probably the single greatest factor in convincing the space agency that single stage to orbit flight was feasible and practical, to borrow from the title of Ivan Bekey's paper of the same name.

    A LO2/kerosene alternative

    A number of people have been asserting for some time that higher propellant mass fractions available from dense propellants may make single stage to orbit possible with those propellants also. The historical examples of the extraordinary mass fractions of the Titan II first stage, the Atlas, and the Saturn first stage are all persuasive. Further, denser propellants lead to higher engine thrust to weight ratios, for perfectly understandable hydraulic reasons.

    It has not usually been observed that higher density also leads to significant reductions in required delta-v. There are two major reasons that this is so. First, the reduction in volume leads to a smaller frontal area and lower drag losses. The second, and more significant, reason is that the gravity losses are also reduced. This is because the mass of the vehicle declines more rapidly from its initial value. The gravity losses are proportional to the mass of the vehicle at any given time, and hence the vehicle reaches its limit acceleration speed faster.

    NASA itself has implicitly recognized this effect. When the Access to Space Option 3 team examined tripropellant vehicles, the delta-v to orbit derived from their work was 29,127 ft/sec, for precisely the reasons described in the previous paragraph. This compares to a delta-v of 30,146 ft/s for the hydrogen-only baseline, as reported in a briefing by David Anderson of NASA MSFC dated 6 October 1993. To be clear, these delta-v numbers include the back pressure losses, so that no "trajectory averaged Isp" number is used. They did not, however, report any results for kerosene-only configurations.

    To come to a more thorough understanding of the issues involved in SSTO design, I have used the same methodology as the Access to Space team to develop compatible numbers for a LO2/kerosene SSTO. There are four major changes in basic assumption between the two approaches, which I will identify and justify here:

    1: The ascent delta-v for the LO2/kerosene vehicle is 29,100 ft/sec, rather than 29,970 ft/sec. The reason for this is argued above, but I ran POST to verify this value, just to be sure. The target orbit is the same: 220 n.mi. circular at 51.6 degrees inclination. The detailed weights I have for the NASA vehicle are based on a delta-v of 29,970 ft/sec rather than the 30,146 ft/sec reported in Anderson's work, but I prefer to use the values more favourable to the hydrogen case to be conservative. The optimum value of thrust to weight ratio turns out to be slightly less than the hydrogen vehicle: 1.15 instead of 1.20.

    2: The aerodynamic configuration is that of Boeing's RASV. Without arguing whether this is optimal, the fineness ratio of 8.27 and large wing lead to a much more airplane-like layout, better glide and crossrange performance, and reduced risk. The single vertical tail is simpler and safer than winglets as well. Extensive analysis has justified the reentry characterisitics of this aircraft. The wing is assumed to be wet with the kerosene fuel, as is common on
    most aircraft. The fuel is also present in the wing carry-through box. The payload is carried over the wing box, and the oxidizer tank is over the wing. This avoids the need for an intertank, which in the NASA Access to Space design is nearly 6,600 pounds.

    3. The main propulsion system is the NK-33. The engine has a sea level thrust of 339,416 lbs, a weight of 2,725 lbs with gimbal, and a vacuum Isp of 331 seconds. Furthermore, it requires a kerosene inlet pressure of only 2 psi absolute, which dramatically reduces the pressure required in the wing tank. It also operates with a LO2 pressure at the inlet of only 32 psi. The comparable values for the SSME are about 50 psi for both propellants. This will have a substantial effect on the pressurization system weight.

    4. The OMS weight is based on the D-58 engine. This engine is used for the Buran OMS system and the Proton stage 4. As heavy as it is the Isp is an impressive 354 seconds. NASA's vehicle used a pressure fed OMS, which is a sensible design choice if you're stuck with hydrogen and you wish to minimize the number of fluids aboard the vehicle. But because both oxygen and kerosene are space-storable, there is no reason to burden the design with a heavy pressure fed system.

    Using the same methodology for calculating masses, and accepting the subsystems masses as given in the Access to Space vehicle, a redesign with oxygen and kerosene was accomplished. The results appear in Table 1.

    Table 1: Access to Space vehicle and LO2/kerosene alternative

    Name                           O2/H2      LO2/RP
    Wing                          11,465      11,893 lb
    Tail                           1,577       1,636 lb
    Body                          64,748      33,741 lb
                Fuel tank         30,668           - lb
              Oxygen tank         13,273      17,271 lb
          Basic Structure         14,610      10,274 lb
      Secondary Structure          6,197       6,197 lb
    Thermal Protection            31,098      21,238 lb
    Undercarriage, aux. sys        7,548       5,097 lb
    Propulsion, Main              63,634      36,426 lb
    Propulsion, RCS                3,627       1,234 lb
    Propulsion, OMS                2,280         823 lb
    Prime Power                    2,339       2,339 lb
    Power conversion & dist.       5,830       5,830 lb
    Control Surface Actuation      1,549       1,549 lb
    Avionics                       1,314       1,314 lb
    Environmental Control          2,457       2,457 lb
    Margin                        23,116      16,105 lb
    Empty Weight                 222,582     141,682 lb

    Payload                       25,000      25,000 lb

    Residual Fluids                2,264       1,911 lb
         OMS and RCS               1,614       1,261 lb
          Subsystems                 650         650 lb
    Reserves                       7,215       8,895 lb
         Ascent                    5,699       7,587 lb
            OMS                      679         541 lb
            RCS                      837         767 lb
    Inflight losses               13,254      17,445 lb
             Ascent Residuals     10,984      15,175 lb
          Fuel Cell Reactants      1,612       1,612 lb
      Evaporator water supply        658         658 lb
    Propellant, main           2,054,612   3,034,972 lb
         Fuel                    293,604     843,048 lb
         Oxygen                1,761,008   2,191,924 lb
    Propellant, RCS                2,814       2,556 lb
         Orbital                   2,051       1,756 lb
         Entry                       763         800 lb
    Propellant, OMS               19,357      15,452 lb
    GLOW                       2,347,098   3,246,156 lb
    Inserted Weight              292,486     211,185 lb
    Pre-OMS weight               271,482     186,152 lb
    Pre-entry Weight             252,125     170,700 lb
    Landed Weight                251,362     169,900 lb
    Empty weight                 222,582     141,682 lb

    Sea Level Thrust           2,816,518   3,733,080 lb
    Percent margin                 11.6%       12.8%   
    Assumed Isp(vac)               447.3       331.0 s
    Ascent Delta-V                29,970      29,100 ft/s
    OMS delta-V                    1,065         987 ft/s
    RCS delta-V                      108         107 ft/s
    Deorbit Delta-V                   44          53 ft/s
    Reserves                       0.28%       0.25% lb/lb
    Residuals                      0.53%       0.50% lb/lb
    Wing Parameter                 4.56%       7.00% lb/lb
    TPS parameter                 12.37%      12.50% lb/lb
    Undercarriage parameter        3.00%       3.00% lb/lb
    Wing Reference Area            4,189       5,528 ft2
    Density of fuel                  4.4        50.5 lb/ft3
    Density of oxygen               71.2        71.2 lb/ft3
    Volume of fuel                66,276      16,694 ft3
    Volume of oxygen              24,733      30,785 ft3
    Fuel tank parameter             0.42           - lb/ft3
    Oxygen tank parameter           0.48        0.56 lb/ft3

    Some discussion of the results and justification is in order.

    The wing is about 40 percent heavier as a percentage of landed weight than for the hydrogen fueled baseline. When considered as a tank, it is about 60 percent heavier for the volume of fuel it encloses. Its weight per exposed area is about the same and the wing loading is half at landing. No benefit is taken explicitly for the lack of a requirement for kerosene tank cryogenic insulation.

    The tail is assumed to have the same proportion of wing weight for both cases. This is conservative for the kerosene wehicle because its single vertical tail is structurally more efficient.

    The body of the kerosene vehicle has three components. The oxidizer tank has an increased unit weight of about 17 percent. This is done in order to avoid the need for aluminum-lithium, which was assumed in the Access to Space vehicle. The basic structure group is unchanged, except that the intertank is deleted and the thrust structure is increased in proportion to the change in thrust level. The secondary structure group is mostly payload support related, and was not changed.

    The thermal protection group is in both cases about 12.5% of the entry weight. This works out to 1.107 lbs/ft2 of wetted area for the kerosene vehicle, which is common to many SSTO designs.

    The undercarriage group is 3% of landed weight for both vehicles. There is no benefit taken for reductions in gear loads for the kerosene vehicle due to lower landing speed and lower glide angle at landing.

    The main propulsion group includes engines, base mounted heat shield, and pressurization/feed weights. The engines are far lighter for their thrust than SSME derivatives. The pressurization weights are reduced in proportion to the pressurized volume for the kerosene vehicle. No benefit is taken for reduced tank pressure.

    Here is as good a place as any to point out the erroneous assertion that increased hydrostatic pressure is going to lead to increased tankage weights. There is no requirement for a particular ullage pressure except for the need to keep the propellants liquid. It is the pressure at the base of the fluid column rather than the top of the column that is of engineering interest. The column of fluid exerts a hydrostatic load on the base of the tank, but this load does not typically exceed the much more adverse requirement for engine inlet pressurization. For the kerosene vehicle, the hydrostatic load at the base of the oxygen tank is 49 psi, which is compatible with the pressures normally seen in oxygen tanks for rocket use. The load declines after launch because the weight goes down faster than the acceleration goes up.

    The bottom line here is that dense propellants may require you to alter a tank's pressurization schedule, but not to overdesign the entire tank. Structures are sized by loads and tankage for rockets is sized principally by volume, and if the vehicle is small, by minimum gauge considerations. This is not completely true for wet wings, however, as discussed previously. In this particular example, there is no need for high pressure in the wing tank either, because of the low inlet pressure required by the NK-33.
     
    The OMS group is the only other major change, as discussed above. The reliable D-58 engine has been performing space starts for decades and will serve well here. The acceleration available from the OMS is about 0.12 g, which is standard.

    All the other weights are pushed straight across for the most part. A brief inspection suggests that this is very conservative. Control surface actuation requirements are certainly less, electrical power requirements less, much better fuel cells available than the phosporic acid type assumed here, and reduced need for environmental control. Nonetheless, rather than dispute any of these values it is easier simply to accept them.

    The margin is applied to all weight items at 15% execpt for the engine group at 7.5%. The justification for this is that the main and OMS engine weights are known to high accuracy.

    The vehicle has an overall length of 1955 inches, and a diameter of 236.4 inches. The wing has a leading edge sweep of 55.5 degrees and a trailing edge sweep of -4.5 degrees. Its reference area is 5,632 square feet, of which 3,992 square feet is exposed. The wing encloses 16,694 ft3 of fuel, with a further 5% ullage. The carry-through is also wet with fuel. The wing span is 1293 inches, and the taper ratio is 0.13.

    The payload bay has a maximum width and height of 15 feet. It sits on top of the wing carry through box. The thrust structure from the engines passes through and around the payload bay to the forward LO2 tank. The payload bay is 30 feet in length. It has a pair of doors, the aft edge of which is just forward of the vertical tail leading edge.

    The engine section encloses 11 NK-33 engines, with a 4 - 3 - 4 layout. The engines are each 12.5 feet long, and additional structure and subsystems take up another 6.5 feet.

    The oxygen tank comprises the forward fuselage, which encloses 30,785 ft3 of oxygen, with a further 5% ullage. The length of the tank is about 100 feet. The ventral surface of the tank is moderately flattened as it moves aft, to fair smoothly with the wing lower surface. This flattening reduces its length by about 5% with respect to a strictly cylindrical layout. The aft edge of the oxygen tank is about even with the forward payload bay bulkhead. A compartment of about 13.9 feet provides room for some subsystems and a potential cockpit in future versions.

    Conclusion

    The methods of the NASA Access to Space study were used to design a single stage to orbit vehicle using existing LO2/kerosene engines. An inspection of the final results shows that the vehicle weighs about 36.5% less than its hydrogen counterpart, with reductions in required technology level and off the shelf engines. The center of mass of the vehicle is about 61% of body length rather than 68% for the Access to Space vehicle, which should improve control during reentry. The landing safety is considerably improved by lower landing speed and better glide ratio. Structural margins are greater overall. The vehicle designed here appears to be superior in every respect: smaller, lighter, lower required technology, improved safety, and almost certainly lower development and operations cost.


    EDIT:  Corrected Roton fuel, per HMXHMX's  post, below.
    « Last Edit: 04/28/2016 08:06 am by Proponent »

    Offline Proponent

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    Re: RP-1, methane, impulse density
    « Reply #108 on: 04/27/2016 03:48 pm »
    Atlas, the classic example of a balloon tank, was 34kPa:

    https://en.wikipedia.org/wiki/SM-65_Atlas

    That pressure was used to stabilize dry Atlases on the ground.  Fueled vehicles in flight surely required much higher pressures.

    Quote
    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.

    There's a difference between stage mass and tank mass, as Whitehead discusses.  He presents real-world tank masses and finds that they scale approximately with volume over two or three orders of magnitude, though the constant of proportionality is noticeably higher when hydrogen is involved.

    You can also have a look at Figure 3 on p. 43 of the attached report on scaling laws.  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.

    Mitch-Burnside Clapp once put together some data on stage masses without engines.  When I have time I'd like to look into that and possibly build a model based on those data, but that's a project for another day.

    Offline Rei

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    Re: RP-1, methane, impulse density
    « Reply #109 on: 04/27/2016 04:43 pm »
    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.

    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.

    Quote
    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.

    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 part by Tom Clancy.  It's pretty hard to take seriously.  It's like using Armadillo Aerospace as a reference.

    The most serious - in terms of buy-in by experts - attempts thusfar were the DC-X and the X-33.  DC-X was ultimately of questionable scalability, and X-33 was scalable but too much trouble to actually get to work with the materials technology available.

    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 masses

    That'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?

    Interesting though that he makes my above point:

    It is the pressure at the base of the fluid column rather than the top of the column that is of engineering interest. The column of fluid exerts a hydrostatic load on the base of the tank, but this load does not typically exceed the much more adverse requirement for engine inlet pressurization.

    Quote
    There's a difference between stage mass and tank mass, as Whitehead discusses.

    By arguing for a linear relationship of tank mass / tank volume, you're arguing for a 3x worse mass ratio for LOX/LH than LOX/kerosene (which, as has been well covered, is not backed up by reality).  Are you trying to argue that LOX/LH engines perform less than one third as well per kilogram of engine mass vs. LOX/Kerosene?  Because if not then counting the full stage mass (rather than just tank mass) will only improve the ratio.

    As I wrote about Whitehead: it's not peer reviewed and he's arguing based overwhelmingly on the Shuttle, which is by no means a "normal" rocket in any stretch.   In fact, it's pretty much the worst rocket in existence you could pick for mass estimation purposes.

    Quote
    You'll see that once the mass of the "propellant module" exceeds 1000 lbm, propellant mass is quite proportional to module mass.

    Propellant mass and propellant volumes are very different things.  My complaint is the concept of scaling tank mass proportional to propellant volume.  Scaling by propellant mass is at least closer to real-world rockets, although still a significant oversimplification.

    Quote
      It's also the case that propellant mass is inversely proportional to to tank pressure, which is consistent with pressure-vessel principles.

    As I - and now the Burnside link that you yourself posted - pointed out, ullage pressure is not as significant as simply the weight of the propellant undergoing the forces of launch.  One doesn't describe a tank whose main job is just to withstand the force of the liquid within it against gravitational/acceleration forces "a pressure vessel".
    « Last Edit: 04/28/2016 01:59 pm by Rei »

    Offline Rei

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    Re: RP-1, methane, impulse density
    « Reply #110 on: 04/27/2016 05:06 pm »
    Hmm, let's just run a couple more datapoints that I haven't yet run -  SLS:

    Stage 1: Dry mass 85200kg, propellant mass 894182kg, using a density of 0,3g/cm = 2,981m^3 = 28,6kg/m³
    Stage 2: Dry mass 3490kg, propellant mass 27220kg, same density = 90,7 m³ = 38,5kg/m³

    And that's not even lithium-aluminum (assuming the mass figures on Wikipedia are up to date)

    Again: real world systems aren't even close to linear scaling with volume.  They're much closer to linear scaling with mass.  SpaceX makes some of the best ratio hydrocarbon rockets in the world, and even they don't approach even poor LOX/LH rockets (with the exception of the Ariane upper stage, due to extenuating factors) in terms of tank kg/m³.  And many of the techniques SpaceX uses to help keep their rockets light could be applied to LOX/LH as well.
    « Last Edit: 04/27/2016 05:11 pm by Rei »

    Offline RanulfC

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    Re: RP-1, methane, impulse density
    « Reply #111 on: 04/27/2016 06:12 pm »
    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 masses

    That'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. The exact study and options are cited and the changes are noted and the math provided:
    "Four major changes were made in assumptions. First, the aerodynamic configuration was changed from a wing with winglets to a swept wing with vertical tail. The delta-V for ascent was as a result recalculated, yielding a lower value due to different values for drag and gravity losses. The engines were changed to LO2/kerosene burning NK-33 engines, which have a much lower Isp than SSME-type engines used in the access to space study, but also have a much higher thrust-to-weight ratio. The orbital maneuvering system on the Access to Space Vehicle was replaced with a pump-fed system based on the D-58 engine used for that purpose now on Proton stage 4 and Buran. Finally, the wing of the vehicle was allowed to be wet with fuel, which is a reasonable practice with kerosene but more controversial with oxygen or hydrogen. Additionally, in order to reduce the technology development needed, the unit weights of the tankage were allowed to increase by 17 percent.

    After the design was closed and all the weights recalculated, the empty weight of the LO2/kerosene vehicle was 35.6% lighter than its hydrogen fuelled counterpart."


    The only reason the number are "out of context" is because you didn't refer to the cited source. *= The author in fact cited the changes made and the calculated effect of those changes and states so right in the beginning. Hint; Don't "skim" and comment as you'll probably be working from wrong information and assumptions.

    Oh and as for "peer-reviewed" as a criteria both Whitehead and Burnsides work was submitted to organizations such as AAIA and NASA for publication and review and both organizations have copies on file and available for reading. I can understand why you might not consider that "peer-review" but stop to consider 99% of the "studies/papers" for things like the DCX and X-33 program were never "peer-reviewed" in the standard sense outside such organizations either.

    Randy
    From The Amazing Catstronaut on the Black Arrow LV:
    British physics, old chap. It's undignified to belch flames and effluvia all over the pad, what. A true gentlemen's orbital conveyance lifts itself into the air unostentatiously, with the minimum of spectacle and a modicum of grace. Not like our American cousins' launch vehicles, eh?

    Offline Rei

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    Re: RP-1, methane, impulse density
    « Reply #112 on: 04/27/2016 08:44 pm »
    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 masses

    That'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.


    "Proof by ghost reference" means "it's not in the reference cited".  I see no table that has the fields he lists.  I find none of his numbers in the document.  Nor do searches for terms he uses like "basic structure" or "undercarriage" yield anything.  If you're seeing them somewhere, I have no clue where.

    Furthermore, the document does consider hydrocarbon options.  Repeatedly.
    « Last Edit: 04/27/2016 08:48 pm by Rei »

    Offline Robotbeat

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    Re: RP-1, methane, impulse density
    « Reply #113 on: 04/27/2016 09:40 pm »
    ....
    All serious efforts toward SSTOs have used hydrogen.....
    ...which may be a very large reason none have ever got off the ground.

    The exception would be the original Atlas. The first US ICBM, the first US orbital crewed rocket, and it was an SSTO, minus the jettisoned side rocket engines (if you replaced all of the engines with an engine of either NK-33's or Merlin 1D's performance, shedding the engines would be unnecessary). Atlas didn't use hydrogen.

    ...and I don't think pushing for a /reusable/ SSTO would improve the situation. If anything, it makes the mass fraction problem worse. It also gives you a much larger vehicle (by volume), which complicates everything.

    Hydrogen is great for an upper stage. Not sure it's great for a SSTO RLV. Hydrogen tends to move you towards drop-tanks due to the high tank mass. Even interplanetary NTR stages tend to rely on drop tanks for this reason. Without shedding tankage, the parasitic mass of those ginormous hydrogen tanks (with their extensive insulation) tends to basically kill any advantage you might have had from hydrogen's high Isp.

    Using a dense fuel means you don't need to shed tank mass, and so you can have a true SSTO, no drop tanks like Shuttle or Rhombus, etc.


    If you DO decide to use hydrogen, you should operate with some way to improve thrust at the beginning, such as injecting LOx or a third propellant in the bell nozzle or something. Also, slush hydrogen starts to look interesting or even necessary.
    « Last Edit: 04/27/2016 09:46 pm by Robotbeat »
    Chris  Whoever loves correction loves knowledge, but he who hates reproof is stupid.

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    Offline HMXHMX

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    Re: RP-1, methane, impulse density
    « Reply #114 on: 04/27/2016 10:36 pm »
    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.

    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.

    Quote
    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.

    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...



    A few corrections, merely to set the record straight and not relevant to the tank mass discussion; while Tom Clancy invested $1M in Rotary, the company raised approximately $30M, so your characterization of him funding it "in large part" is in error. And while the demonstrator was hard to fly, the point of conducting the ATV flight program was to identify and correct any control deficiencies in future vehicles, which would have been done if the program had continued, of course.  As for TPS, we indeed did have a heat shield in development that worked perfectly well.  Finally, the only "scrap parts" in the vehicle were the rotor hub and blades repurposed from a certified helicopter.  Calling them scrap (they were FAA certified) is pejorative and inaccurate.

    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.  So I won't argue that point with you.

    But for other readers edification, I'll provide the mass information for the Roton C9 airframe:

    LOX Tank Mass: 1049.4 lbm (as built) for 15 psia ullage
    JP4 Tank Mass:    242.7 lbm (as built) for 15 psia ullage
    LOX Prop Load:  315,364 lbm
    JP4 Prop Load:    83,000 lbm

    Vehicle Landed Mass: 22,446.7 lbm (as built and calculated from drawings)
    Mass Ratio: 15.05 (crew mission only)
    PMF: 0.9335

    Tank mass was less than 0.5% of propellant mass.  Naturally, because of the vehicle configuration, this doesn't include the large intertank structure, nor engine thrust structure, but is illustrative of what can be accomplished with good design and attention to ullage pressure specifications.



    « Last Edit: 04/27/2016 10:37 pm by HMXHMX »

    Offline jongoff

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    Re: RP-1, methane, impulse density
    « Reply #115 on: 04/28/2016 06:08 am »
    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.htm

    Hydrogen 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.

    Yeah, the only three places where LOX/LH2 makes reasonable sense are IMO:

    1- In-space stages (since every kg of IMLEO is going to be far more expensive than a kg on the ground)
    2- Upper stages (since it can dramatically reduce the size of the first stage for a given payload, if done properly)
    3- Air-launched TSTOs or SSTOs (there are some real hassles with making LH2 work for air-launch, but air-launched rockets are pretty much the only rockets that are strongly GTOW constrained).

    Trying to use LOX/LH2 for ground-launched first stages or TSTOs seems like a really bad idea.*

    ~Jon

    * Unless you're using Thrust Augmented Nozzles with the nozzle section running very lean of stoichiometic. That might plausibly allow a LOX/LH2 booster stage or SSTO to make more sense, I haven't run enough numbers to convince myself one way or another.

    Offline Proponent

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    Re: RP-1, methane, impulse density
    « Reply #116 on: 04/28/2016 08:12 am »
    LOX Prop Load:  315,364 lbm
    JP4 Prop Load:    83,000 lbm

    Since 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?
    « Last Edit: 04/28/2016 08:13 am by Proponent »

    Offline Rei

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    Re: RP-1, methane, impulse density
    « Reply #117 on: 04/28/2016 11:01 am »
    $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.

    Don't get me wrong - I respect that you were trying something new, and I respect that you were trying to get it done with a minimal budget and had to go Kerbal to some extent to do it.  But when people call that "the most serious SSTO effort ever"... obviously these sorts of things are going to come up.

    Quote
    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.

    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).  Demonstrably

    Asserting 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.)

    Quote
    LOX Prop Load:  315,364 lbm
    JP4 Prop Load:    83,000 lbm
    Vehicle 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.
    « Last Edit: 04/28/2016 02:11 pm by Rei »

    Offline HMXHMX

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    Re: RP-1, methane, impulse density
    « Reply #118 on: 04/28/2016 02:06 pm »
    LOX Prop Load:  315,364 lbm
    JP4 Prop Load:    83,000 lbm

    Since 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?

    It was pretty obvious well before we began detail design that hydrocarbons were the way to go, based on simple analysis of PMF efficiency as a function of Density ISP, but also the desire to have dense propellants to deliver high pressure in the chambers (unique to Roton which employed centrifugal pumping of propellants by rotating the whole engine).  JP4 was the simplest solution at the time, and probably remains so for this type of design.

    Pressurization was still being traded when we shut down, but baseline was autogenous for LOX and warm N2 for JP4, as I recall. 
    « Last Edit: 04/28/2016 02:14 pm by HMXHMX »

    Offline HMXHMX

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    Re: RP-1, methane, impulse density
    « Reply #119 on: 04/28/2016 02:11 pm »
    $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.

    Quote
    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.

    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).  Demonstrably

    Asserting 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.)

    Quote
    LOX Prop Load:  315,364 lbm
    JP4 Prop Load:    83,000 lbm
    Vehicle 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.


    All the components of ATV whose weight I quoted were built.  The ATV had its JP4 tank replaced by the H2O2 tank and pressurization system, but otherwise was the same structure as the PTV (Propulsion Test Vehicle).

    I'll stop trying to educate you on your other snarky comments; life is to short.

    Offline Rei

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    Re: RP-1, methane, impulse density
    « Reply #120 on: 04/28/2016 02:41 pm »
    Showing that the very numbers you quoted are literally double the mass/propellant volume of modern hydrogen stages is not "snark", it's the very point of this whole discussion: whether the concept of some propellant-ambivalent mass/volume ratio is backed up by real-world hardware.

    And one can't just "replace a tank" in a rocket and call it the same thing, or ascribe numbers on a vehicle lacking a rocket engine and TPS that is only run thrhough a very minimal flight envelope (and can hardly hold itself stable in that) as if they're numbers on some full, functional vehicle.  You're presenting estimates as if they're hard stage numbers.  And even those estimates aren't impressive compared to modern hydrogen stage in terms of mass/propellant volume ratios - literally half as good.

    Which is, again, the whole point of this conversation.

    If one is running estimates about what makes the best propellant combination based on a mass premise that is wrong by twofold in an optimistic case, and half an order of magnitude more often than not in the real world, then those estimates are going to be dramatically, dramatically wrong.
    « Last Edit: 04/28/2016 02:48 pm by Rei »

    Offline Chris Bergin

    Re: RP-1, methane, impulse density Q&A
    « Reply #121 on: 04/28/2016 02:52 pm »
    Thread is done. Several people have tried to help one member, but that member isn't having it.

    Locked.
    « Last Edit: 04/28/2016 03:09 pm by Chris Bergin »
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