Quote from: hkultala on 04/14/2016 09:01 pmGood point. For perfect burning, mixture ratio should be 16.You mean 8, right?
Good point. For perfect burning, mixture ratio should be 16.
QuoteIt would seem that RS-68 uses too small mixture ratio;Maybe it's partly to increase the thrust-to-weight ratio.
It would seem that RS-68 uses too small mixture ratio;
Hmm, a thought just occurred to me, wherein cryochilled propane's viscosity could be an advantage. Are you familiar with the research on metalized gel propellants? The concept is to add gelling agents like fumed silica to allow you to suspend metal dusts like aluminum powder in the liquid rocket fuel. Aluminum combustion gives off a great deal of energy for its mass, providing additional heat to the exhaust stream. But you know, gelling basically means "increasing the viscosity". With cryochilled propane, you already have increased viscosity vs. "runny" fuels like RP-1 (whether it's sufficient to suspend aluminum particles without gelling agents, that I can't say).
To further develop my understanding of the trade between specific impulse and density, I've done a little thought experiment on ground-launch stages.
1. Except for JP-5 (the composition of which I don't know),
2) Some cooling-related thing makes the engine T/W not scale with propellant density. The extra unburned H2 makes the engine run cooler and allow less mass to be used for cooling?
I've never seen any indication of a 'viscosity' issue with cryo-propane. It's denser than when liquid under normal pressure/temperature but nothing that impedes either turbo-pump or pressure fed use. When tested in the RL10 it was less 'viscos' than RP1 and more like LH2 which was a cited advantage in that type of engine.
So if I read all this right then I get that balancing both impulse and density has not been as straight-forward as even the experts (I'm thinking all the early "when we have hydrogen we can do anything" rocket scientist here ) had thought. Further it would seem that in a TSTO system it might be more efficient to consider different propellants for booster and upper stage despite a slightly higher operations costs?
JP-5 or JP-10? JP-5 Material Safety Data Sheets, (MSDS) are available on-line and composition properties sheets IIRC ....
Quote from: Impaler on 06/09/2015 03:22 amSteven Pietrobon: I think Zubrin himself has fully acknowledged that Ethylene is completely superior to Methane and if he had the whole thing to do over again he would have pushed that instead as it's synthesis is almost as easy as methane, higher hydrocarbons not so much.Lower hydrogen needs for Ethylene and easier refrigeration (practically none on Mars) are considered even more important then the density and impulse values. The only reason to go for Methane now is that fact that everyone is developing LNG based engines for launch vehicles now and you could reuse thouse engines on Mars, but even then I suspect a dual fuel engine would be possible and advantageous.Refrigeration advantages are moot when sharing a thermal environment with LOx. Ethylene is a moderate problem there, because to maintain it in liquid phase at the same temperature you would need to raise LOx tank pressure to 5+ atmospheres (ethylene freezes at the boiling point of LOx at about 3.5atm). That's manageable, but adds weight. Zubrin is currently working on ethylene-N2O green hypergolics - http://www.parabolicarc.com/2015/05/15/pioneer-astronautics/
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.
I've been told on several occasions that Lox-Methane is a very feasible mixture for a Mars mission, with references to the Morpheus lander, though I was curious about its ignition. The advantages I saw listed were non-toxicity, low cost of propellant, and lower energy use versus hypergolic propellants. The obvious disadvantages were impulse density and ignition issues. Aside from gaining hypergolic ignition with an Ethylene-Nitrous Oxide mix, how does its compare with the obvious Lox-Methane and Lox-Ethylene alternatives in terms of advantages and disadvantages?
Quote from: Hyperion5 on 04/16/2016 01:10 amI've been told on several occasions that Lox-Methane is a very feasible mixture for a Mars mission, with references to the Morpheus lander, though I was curious about its ignition. The advantages I saw listed were non-toxicity, low cost of propellant, and lower energy use versus hypergolic propellants. The obvious disadvantages were impulse density and ignition issues. Aside from gaining hypergolic ignition with an Ethylene-Nitrous Oxide mix, how does its compare with the obvious Lox-Methane and Lox-Ethylene alternatives in terms of advantages and disadvantages?Attached are plots showing the performance of both oxygen and nitrous oxide with light hydrocarbons. Low-orbit speed on Mars is about 3.5 km/s. Since the atmosphere is thing and the gravity weak, perhaps 4 km/s is not too optimistic as a delta-V for getting from the surface to low orbit. The first plot shows that both oxygen and nitrous do pretty well. Oxygen is better, but nitrous isn't too bad.On the other hand, the thin atmosphere means that a martian SSTO making a round trip would have a pretty large delta-V to perform in returning to the surface. Taking an approximate, worst case, suppose the total delta-V to orbit and back is 8 km/s. Then, as the second plot shows, the performance of nitrous is really rather poor.
... the thrust should be about linearily propotional to the propellant density ....
I'm interested in seeing plain N2O myself. I always wondered at why it was not used given its ease of storage, self pressurization and ability to be A) a monopropellant and B) a ignition source with a reusable catalyst. It's performance really isn't THAT bad, but in a reusable system that isn't 100% performance optimized I always wondered why it didn't never really saw much consideration.
through my amateurish calculations expect a high 240s-range sea-level ISP
******************************************************************************* NASA-GLENN CHEMICAL EQUILIBRIUM PROGRAM CEA2, MAY 21, 2004 BY BONNIE MCBRIDE AND SANFORD GORDON REFS: NASA RP-1311, PART I, 1994 AND NASA RP-1311, PART II, 1996 ******************************************************************************* prob rocket fac p,bar=200.0 ions pi/pe=1276.851685 mdot=2223.8 reac fuel=N2O moles=1.0 t(k)=300.0 outp short end THEORETICAL ROCKET PERFORMANCE ASSUMING EQUILIBRIUM COMPOSITION DURING EXPANSION FROM FINITE AREA COMBUSTOR Pin = 2900.8 PSIA MDOT/Ac = 2223.800 (KG/S)/M**2 Pinj/Pinf = 1.003364 CASE = REACTANT MOLES ENERGY TEMP KJ/KG-MOL K FUEL N2O 1.0000000 81671.539 300.000 O/F= 0.00000 %FUEL=100.000000 R,EQ.RATIO= 0.000000 PHI,EQ.RATIO= 0.000000 INJECTOR COMB END THROAT EXIT Pinj/P 1.0000 1.0068 1.8334 1276.85 P, BAR 200.00 198.66 109.09 0.15664 T, K 1908.42 1907.02 1670.30 305.99 RHO, KG/CU M 3.6988 1 3.6766 1 2.3049 1 1.8065-1 H, KJ/KG 1855.63 1853.80 1550.33 7.8034 U, KJ/KG 1314.91 1313.48 1077.06 -78.903 G, KJ/KG -12294.6 -12287.8 -10835.9 -2261.27 S, KJ/(KG)(K) 7.4146 7.4156 7.4156 7.4156 M, (1/n) 29.345 29.345 29.344 29.342 (dLV/dLP)t -1.00007 -1.00007 -1.00003 -1.00000 (dLV/dLT)p 0.9998 0.9998 0.9998 1.0000 Cp, KJ/(KG)(K) 1.3049 1.3046 1.2594 0.9960 GAMMAs 1.2771 1.2772 1.2901 1.3976 SON VEL,M/SEC 831.0 830.7 781.4 348.1 MACH NUMBER 0.000 0.073 1.000 5.522 PERFORMANCE PARAMETERS Ae/At 8.0991 1.0000 51.862 CSTAR, M/SEC 1106.7 1106.7 1106.7 CF 0.0547 0.7060 1.7370 Ivac, M/SEC 8993.6 1387.1 1967.5 Isp, M/SEC 60.5 781.4 1922.4 MOLE FRACTIONS *NO 0.00676 0.00673 0.00299 0.00000 NO2 0.00024 0.00024 0.00013 0.00000 *N2 0.66324 0.66325 0.66515 0.66667 *O 0.00001 0.00001 0.00000 0.00000 *O2 0.32974 0.32975 0.33173 0.33333 * THERMODYNAMIC PROPERTIES FITTED TO 20000.K
To further develop my understanding of the trade between specific impulse and density, I've done a little thought experiment on ground-launch stages.In 1996, John Whitehead wrote a cute little paper about SSTO mass budgets (4th attachment to this post. For a few different propellant combinations, he used the rocket equation to calculate the mass ratios needed for a delta-V of 10 km/s (i.e., Earth to LEO with losses). Then he estimated the masses of engines, tanks, pressurants and residual propellants (the last two can be larger than you expect) as a fraction of burn-out mass. Let's call the part of the burn-out mass that's not devoted to those four things the available mass (I'm open to suggestions for a better term). Many subsystems will have to be crammed into this so-called available mass: landing gear, if any, avionics, etc., etc. The idea, though, is that the masses of such subsystems will be approximately independent of the propellants chosen. Hence, the vehicle with the highest available mass fraction should have the largest payload fraction as well.The name of the game, then, is to choose the propellant combination that maximizes the available mass.About the same time as Whitehead's paper, Bruce Dunn presented an analysis in a similar spirit (3rd attachment). Although Dunn's assumptions were perhaps a bit more ad hoc, he covered a wider range of propellants.I've made a similar calculation similar to Whitehead's. There are just two major differences. Firstly, Whitehead assumes that the lift-off thrust-to-weight ratio of the engine is a linear function of propellant density and is 100 for lox/RP-1 and 50 for lox/hydrogen. In contrast, I assume the ratio is proportional to the impulse density of the propellants (it seems to me this makes more sense; any comments?), taking a value of 123 for lox/RP-1 at a typical mixture ratio (essentially, the NK-33 or the AJ-26).The second significant difference is that rather than assuming a particular mixture ratio, I adjust the mixture ratio for maximal performance.Otherwise, to oversimplify slightly, I use pretty much the same assumptions: 10 km/s of delta-V, tanks weigh 10 kg/m3, pressurants and residuals are each 0.25% of the initial propellant load. Specific impulses come from RPA Lite 1.2.8 and are scaled by 0.95 from ideal vacuum values. Chamber pressure is 20 MPa and the area expansion ratio is 40:1. For the time being, propellants are assumed to be at the lower of room temperature and the normal boiling points.Have a look at the first plot attached. It shows specific impulses delivered by various fuels1 burned with oxygen as a function of propellant bulk density. Also shown as grey curves are contours of constant "available mass." These contours are easily calculated, since all that's required in Whitehead's model is a specific impulse and a propellant density. The first table, below, gives optimal figures for each of 30 propellant combinations.Hydrogen does poorly. If the mixture ratio is allowed to vary during flight in an optimal way, the available mass fraction with hydrogen as a fuel increases2 by about 0.026. Other fuels don't benefit much from mixture-ratio variation, so the this enough to boost hydrogen to the middle of the table. But, the substantially larger mass of hydrogen tanks arising from the need to insulate them has been neglected. Taking this into account would knock hydrogen right back to the bottom of the table.Speaking of the table, a couple of columns may not be self-explanatory:* Mix: Linear function of the mixture ratio, being zero for maximum Isp and unity for maximum impulse density.* T/W: Thrust-to-weight ratio of the engine at lift-off (giving the a ratio of 1.3 for the vehicle).* Den exp: slope of the log Isp-log(bulk density) curve at the optimum; shows the relative importance of density compared to Isp.People often obsess about maximizing specific impulse. The Mix column shows that's not generally what you want to do.The "Den exp" column shows the relative sensitivity of available mass fraction to density as opposed to specific impulse. For the better performing propellant combinations, it's about 0.23, meaning that a the figure of merit is approximately: (specific impulse)(bulk density)0.23for an SSTO. This is, of course, somewhat model dependent, but it happens to be about the same as what I estimated from Dunn's results some time ago.OK, so, what about hydrogen peroxide, with its high density? Please have a look at the second plot. This time I've left hydrogen out so as to make the hydrocarbons more visible. As you easily see, peroxide's density does not raise bulk density enough to make up for its lower specific impulse. Bruce Dunn told us that a long time ago, but I find it educational to see it graphically. I also looked at nitric acid, which is even denser (1510 kg/m3) than peroxide (1460 kg/m3). It, however, suffers from lower specific impulse and lower bulk density than you might expect: the fact that it contains quite a bit of free oxygen means that mixture ratios with nitric acid tend to be low.If we consider a delta-V of just 4 km/s -- see the third plot and second table -- peroxide looks much better. As you'll see from the table, the figure of merit at this delta-V, which could correspond to a first stage or a martian SSTO, is something like: (specific impulse)(bulk density)0.4 ,Finally, consider a very low delta-V, like 40 m/s, as shown in the final plot. In this case, impulse density reigns, and peroxide is the run-away winner. The associated table shows that the figure of merit is very close to (specific impulse)(bulk density) ,i.e., impulse density, which is just what you expect when delta-V is small compared to exhaust velocity. Note, though, that we do have to go to very low delta-V's before impulse density dominates.All of the above is applies to ground-lit stages. For upper stages, mass will be more important, since the stage's propellant must be accelerated by lower stages. Hence, the density exponent in the figure of merit will tend to be smaller.1. Except for JP-5 (the composition of which I don't know), the color of each curve is the number of carbon atoms, modulo 10, in each fuel's principal chemical component (e.g., 1 for methane, 2 for ethane and ethylene) expressed in the resistor color code. Solid lines are used for saturated hydrocarbons (alkanes). The two alkenes, ethylene and proplylene, are shown with dashed lines.2. If a different mixture ratio is allowed for each successive 1% of the total propellant volume, the ratio ranges from 17.8 (633 kg/m3) at lift-off to 5.7 (350 kg/m3) at burn-out. A variable mixture-ratio program helps in two ways. Firstly, it simply helps with the rocket equation by allowing more impulse to be packed in at the beginning, where mass doesn't matter so much, while going for higher specific impulse at later times. Secondly, it increases the lift-off thrust-to-weight ratio of the engine, allowing for a smaller engine.EDIT: Added "bulk" to very-low-delta-V figure of merit.
, tanks weigh 10 kg/m3
pressurants and residuals are each 0.25% of the initial propellant load
Chamber pressure is 20 MPa and the area expansion ratio is 40:1.
Hydrogen does poorly. If the mixture ratio is allowed to vary during flight in an optimal way, the available mass fraction with hydrogen as a fuel increases2 by about 0.026.
But, the substantially larger mass of hydrogen tanks arising from the need to insulate them has been neglected.
* Mix: Linear function of the mixture ratio, being zero for maximum Isp and unity for maximum impulse density.* T/W: Thrust-to-weight ratio of the engine at lift-off (giving the a ratio of 1.3 for the vehicle).* Den exp: slope of the log Isp-log(bulk density) curve at the optimum; shows the relative importance of density compared to Isp.
(specific impulse)(bulk density)0.23
Quote, tanks weigh 10 kg/m3Can this concept be defended? Mass loadings on the tanks will certainly be different with different propellants at the very least. Not to mention the x^3/x^2 volume/surface area scaling issue....
...When you're talking about SSTOs, the picture doesn't change. You just can't use the sort of low-ISP high-thrust stage you'd use with a staged rocket. You still have to use H2 because SSTOs are even more ISP-dependent than staged rockets:https://en.wikipedia.org/wiki/File:SSTO_vs_TSTO_for_LEO_Mission.tifYou simply can't get a plausible structural coefficient with a low ISP propellant mix. It just doesn't work.
The volume/surface area scaling issue does not apply to pressure vessels (unless you run into minimum-gauge issues or decide to use a large amount of insulation). Since rocket tanks are fairly well approximated as pressure vessels
If you get really, REALLY tall (like Saturn V first stage size), then you have to start taking into account pressure head (and this can actually allow you to SAVE weight, since you can use a little less ullage pressure and the top of the stage can thus be made a little thinner), but for our purposes here, that's a pretty good estimate.
This last one is perhaps the biggest reason why pump-fed rocket engines are used instead of pressure-fed.
Rei, have you read Dunn's report on various SSTO propellant combinations? It is not kind to hydrogen.http://web.archive.org/web/20120303152352/http://www.dunnspace.com/alternate_ssto_propellants.htmHydrogen may have the best Isp, but liquid hydrogen is, in fact, the least dense liquid known to humankind. It has been worshipped by aerospace since Tsiolkovsky, but in no way is it an optimal fuel for a SSTO rocket, particularly a reusable one (where dry mass is yet more important). Please re-examine your prejudices in light of that Dunn report.
In the current model, most propellant combinations beat hydrogen/oxygen. This is a direct result of assuming a constant-size rather than constant-mass vehicle for all propellants, regardless of density.
Falcon 9 v1.0 was thought to have an ullage pressure of about 50psi, that's more than just "a fraction of an atmosphere overpressure."
Additionally, Saturn V is a poor example because the different stages were built by different entities. Additionally, the first stage is obviously going to be built much different than the other stages due to the lower penalty for high dry mass first stage (with its big ol' fins, etc). You should be comparing pump-fed upper stages to other pump-fed upper stages.