Should launch providers switch to common methalox engines?

[oldAtlas_Eguy]

While MethaLox is great for the Earth-Mars run it is not the best prop for cis-lunar travel when there is a source for HydroLox from asteroid or Lunar surface water. In which case in-space only systems would not be MethaLox but HydroLox or even SEP. It is a mater of which fuel is easily available at lower costs per kg. So Earth to orbit and Mars to orbit systems would work best as MethaLox from a cost standpoint, in-space only systems using MethaLox would be more costly than other choices such as HydroLox or LH2 or Argon (found as trace gas in Lunar regolith and on asteriod material as well) used for SEP.

[Stan-1967]

The oxygen in the atmosphere will quickly burn with the very hot excess methane.

Not necessarily. The exhaust of an upper stage rocket engine with a high expansion ratio is generally colder than the flammability point of methane, as the expansion in the nozzle converts most of the thermal heat into mean mechanical motion. The methane will only combust if it hits atmosphere at a high enough speed that it becomes compression heated.

If the exhaust velocity is say, 3.6 km/s, then the exhaust will be released at velocity ~zero when the vehicle is traveling at 3.6 km/s halfway through the launch. That portion of the exhaust will make it into the atmosphere without combusting.

The premise that because methane is burned fuel rich will result in methane emmissions into the atmosphere, and therefore methane rockets are problematic is not supportable and is itself, problematic.

1.  How much methane is problematic?  ( bifurcate this into low altitude vs. high altitude it you'd like to)
2.  Does methane being combusted with LOX at pressures likely to be around 200 bar produce methane as a combustion byproduct?    If so, how much?   See question #1

The concern of methane being "problematic" seems like a manufactured scare tactic base on bad logic, i.e if a thing can happen, it will happen, as well as unsupportable science on the combustion process itself.

[Nilof]

While MethaLox is great for the Earth-Mars run it is not the best prop for cis-lunar travel when there is a source for HydroLox from asteroid or Lunar surface water. In which case in-space only systems would not be MethaLox but HydroLox or even SEP. It is a mater of which fuel is easily available at lower costs per kg. So Earth to orbit and Mars to orbit systems would work best as MethaLox from a cost standpoint, in-space only systems using MethaLox would be more costly than other choices such as HydroLox or LH2 or Argon (found as trace gas in Lunar regolith and on asteriod material as well) used for SEP.

Actually, Methalox is generally a better choice than Hydrolox for an in space stage using propellant from lunar ice or carbonaceous chondrites, since both are carbon rich (lunar ice contains more CO/CO2 than H2O). Due to higher density a single methalox stage will have more performance than a hydrolox stage unless total delta-v is over ~12 km/s. The reason why hydrolox is useful for space launch is because of staging, the hydrogen upper stage has less delta-v but also weighs less.

[Nilof]

The oxygen in the atmosphere will quickly burn with the very hot excess methane.

Not necessarily. The exhaust of an upper stage rocket engine with a high expansion ratio is generally colder than the flammability point of methane, as the expansion in the nozzle converts most of the thermal heat into mean mechanical motion. The methane will only combust if it hits atmosphere at a high enough speed that it becomes compression heated.

If the exhaust velocity is say, 3.6 km/s, then the exhaust will be released at velocity ~zero when the vehicle is traveling at 3.6 km/s halfway through the launch. That portion of the exhaust will make it into the atmosphere without combusting.

The premise that because methane is burned fuel rich will result in methane emmissions into the atmosphere, and therefore methane rockets are problematic is not supportable and is itself, problematic.

1.  How much methane is problematic?  ( bifurcate this into low altitude vs. high altitude it you'd like to)
2.  Does methane being combusted with LOX at pressures likely to be around 200 bar produce methane as a combustion byproduct?    If so, how much?   See question #1

The concern of methane being "problematic" seems like a manufactured scare tactic base on bad logic, i.e if a thing can happen, it will happen, as well as unsupportable science on the combustion process itself.

AFAIK it isn't. But if you want to have an accurate number for the quantity of greenhouse gases released by a launch, you have to include all terms. You can't decide that the option that is more convenient for you is necessarily better before you do the analysis.

[Oli]

Due to higher density a single methalox stage will have more performance than a hydrolox stage unless total delta-v is over ~12 km/s.

Based on what?

[Stan-1967]

AFAIK it isn't. But if you want to have an accurate number for the quantity of greenhouse gases released by a launch, you have to include all terms. You can't decide that the option that is more convenient for you is necessarily better before you do the analysis.

It is a given that an accurate accounting of all chemical species in the rocket plume must be performed as part of understanding the environmental impact of space launches, as well as for the rocket designers to validate the combustion models that predict the species in the plume.   That validation is needed to optimize the thrust and ISP performance of the engine. 

However let's not move the goalpost in going from outright declaring methane problematic, to now stating that the analysis must be done, and we can't just decide on a convenient option without doing the analysis. 
1.  How do you know the analysis has not already been done, and that the companies pouring billions of dollars into development have not considered this already?
2.  How do you justify asserting that methane is in fact problematic, when the significant body of knowledge regarding hydrocarbon combustion science says that the products of methane LOX combustion will rapidly move the methane concentration to extinction levels within the reactants?

Of immediate concern regarding regulatory agreements that affect space launch is production of halogens and particulates that will affect ozone depletion.   Methane is vastly preferable than solid rockets that produces large quantities of Cl and particulates in the plume.  OH, which decomposes from H2O, also depletes local ozone concentrations, but a methane fueled rocket is no worse than kerolox, hydrolox or hypergolic fuels in this regard.   Hypergolics are worse than methane/LOX in that they produce significant NO.

The comments on the preference of Methane vs. Hydrogen/LOX & argon for in space propulsion between you and Atlas guy were very informative.  Which one is preferred may well depend on how advanced the outer space economic ecosystem evolves.   Which propellant is most likely to be "common" may well depend on how specialized a space economy becomes.  It may be one fuel for travel to planetary surfaces, and another fuel for deep space and asteroid sized bodies.  Before such a ecosystem is available, the most preferred and common fuel may simply be what is available first.

[Stan-1967]

To swing the back and forth discussion re: methane environmental concerns back to the thread topic, I think there is the start of enough incentive for existing launch providers to move away from hydrolox, kerolox, and solids.  It is arguable that re-use is starting to create significant pressure in this regards. 

Solids are a nice way to give a rocket alternate configurations that improve payload without completely redesigning the core and embarking on a larger engine program.  A certain company that gets lots of attention hear at NSF is pushing the scales of obsolescence against solids by building a reusable core that is sized not to need them, and it is doing that with kerolox.  Hydrolox as a first stage fuel seems to be less preferred in new designs due to cost/complexity, and not due to anything related to the potential of Methane. 

[Oli]

Black carbon emitted into the stratosphere is the biggest issue (100k the radiative forcing of CO2). CO2 from rockets will likely never be a problem. Methane produces a lot less soot than kerosene (5x-10x?), so it should be better.

Of course solids with AP (Ammonium perchlorate) are the worst, since the hydrogen chloride destroys the Ozone layer. With higher flight rates solids would have to switch to ADN (Ammonium dinitramide).

[A_M_Swallow]

{snip}
The comments on the preference of Methane vs. Hydrogen/LOX & argon for in space propulsion between you and Atlas guy were very informative.  Which one is preferred may well depend on how advanced the outer space economic ecosystem evolves.   Which propellant is most likely to be "common" may well depend on how specialized a space economy becomes.  It may be one fuel for travel to planetary surfaces, and another fuel for deep space and asteroid sized bodies.  Before such a ecosystem is available, the most preferred and common fuel may simply be what is available first.

A propellant depot can have several tanks permitting use of more than one fuel.

First stage Earth launch - methane/LOX
Second stage Earth launch - either methane/LOX or hydrogen/LOX
Cargo Earth to Moon and Mars - Argon
Manned Earth to Moon and Mars - methane/LOX
Lunar lander, fuel from Earth - methane/LOX
Lunar lander, ISRU fuel - hydrogen/LOX

[Nilof]

Due to higher density a single methalox stage will have more performance than a hydrolox stage unless total delta-v is over ~12 km/s.

Based on what?

By computing the payload fraction/dry mass fraction(the latter is also closely related to volume) at various delta-v's, even using very optimistic assumptions for LH2 tankage and very conservative assumptions for LCH4 tankage. To accelerate a given payload to a given delta-v, the methane stage will be more compact and have a lower dry mass. I had a fairly thourough post about this with numbers and graphs in a previous thread, which should be a couple years old by now. I can link to it if I find it.

This is much like how the Falcon 9 has more capability than a single-stick Delta IV even though the Delta has both a bigger tank volume and more dry mass. The Delta IV only catches up at high delta-vees.

Hydrolox is good if you are constrained by initial fuel mass(as is the case for a second stage). If you aren't, you want dense propellants for most delta-v's.

As far as the ISRU side is concerned, if you have free CO2/CO the extra fuel quantity you get from taking a given quantity of hydrolox and using the sabatier process to make Methalox is larger than the increased fuel mass needed due to lower Isp.

Hence, changing to methane also significantly increases the number of flights for a fixed quantity of available hydrogen(or power if the hydrogen comes from electrolysing water), especially at low delta-v's. This is the thinking behind Mars direct, but it is also applicable to the lunar poles since LCROSS found out that the ice in dark lunar craters contains more CO than H2O.

This means that there is more commonality between martian and lunar poles ISRU than most people realize. LH2/LOX is only the better choice for lunar ISRU if you do LOX ISRU from regolith only. If lunar ice is used to make fuel, methane tends to win out.

[sdsds]

In answer to the original question I think there is a case that can be made for a "methalox everywhere" architecture. But the original question limited that to both (or all) stages of a launch vehicle. That seems to me a less interesting question.

ISTM "methalox everywhere" works only in the context of in-situ (or in space) propellant production (ISPP). And it presupposes that something with readily available carbon atoms (CO or CO2, for example) will be available everywhere that e.g. readily available hydrogen in the form of H2O is available.

[Oli]

To accelerate a given payload to a given delta-v, the methane stage will be more compact and have a lower dry mass.

If we assume ambitious numbers for both hydrolox (470s isp, 90% prop. mass fraction) and methalox (380s isp, 93% pmf), hydrolox stages have a lower dry mass from 5.6km/s upwards.

[Pipcard]

In answer to the original question I think there is a case that can be made for a "methalox everywhere" architecture. But the original question limited that to both (or all) stages of a launch vehicle. That seems to me a less interesting question.

ISTM "methalox everywhere" works only in the context of in-situ (or in space) propellant production (ISPP). And it presupposes that something with readily available carbon atoms (CO or CO2, for example) will be available everywhere that e.g. readily available hydrogen in the form of H2O is available.

I agree, I should have asked about the implications of a "all-methalox architecture" instead.

[A_M_Swallow]

Lander engines need to throttle back. Would this feature impose too big a mass or Isp ot thrust penalty on upper stages or transfer vehicles?

Using two different designs of engines will increase the GLOW mass because spacestations will have to have two sets of replacement parts.

[Steven Pietrobon]

Here's the output for the Raptor engine from ISP.EXE using best estimates of chamber pressure, mixture ratio, expansion ratio and nominal boiling point (NBP) propellants. CH4 is not even considered as a species in the exhaust! There's a small amount of hydrogen which will immediately burn with the oxygen in the atmosphere when it exits the nozzle.

   PROPELLANT        HF   DENSITY    WEIGHT     MOLES      VOLUME
 O2             -3.1020    1.1490    3.5000     .1094    3.0461
 CH4           -21.3900     .4239    1.0000     .0623    2.3590

  GRAM ATOMS/100 GRAMS
 H    5.5406 O    4.8613 C    1.3852

  ENTHALPY = -37.16838      DENSITY = .833
  CSTAR =  6077.85

                          CHAMBER    THR(SHIFT)    EXH(SHIFT)    EXH(SHIFT)
  PRESSURE (PSIA)        2662.081      1537.667        14.700         5.931
  EPSILON                    .000         1.000        21.841        44.416
  ISP                        .000       123.737       335.236       353.996
  ISP (VACUUM)               .000       232.855       358.019       372.690
  TEMPERATURE(K)         3697.667      3511.070      2123.630      1841.663
  MOLECULAR WEIGHT         22.200        22.483        24.039        24.061
  MOLES GAS/100G            4.505         4.448         4.160         4.156
  CF                         .000          .655         1.775         1.874
  PEAE/M (SECONDS)           .000       109.118        22.783        18.694
  GAMMA                     1.195         1.194         1.194         1.200
  HEAT CAP (CAL)           54.765        54.494        50.982        49.610
  ENTROPY (CAL)           279.835       279.835       279.835       279.834
  ENTHALPY (KCAL)         -37.169       -54.757      -166.272      -181.126
  DENSITY (G/CC)      1.32533E-02   8.16494E-03   1.37987E-04   6.42562E-05
  ITERATIONS                    8             3            10             8
   
  MOLES/100 GRAMS 
   
 H                         .07998        .06778        .00357        .00083
 H2                        .31930        .29805        .23807        .26736
 HO                        .29786        .25063        .00479        .00058
 HO2                       .00130        .00082        .00000        .00000
 H2O                      2.26118       2.31250       2.52806       2.50226
 H2O2                      .00019        .00011        .00000        .00000
 O                         .04244        .03285        .00005        .00000
 O2                        .11716        .09996        .00023        .00001
 HCOOH                     .00003        .00002        .00000        .00000
 CHO                       .00010        .00006        .00000        .00000
 CH2O                      .00001        .00000        .00000        .00000
 CO                        .74768        .70673        .44236        .41183
 CO2                       .63734        .67835        .94279        .97332

[oldAtlas_Eguy]

To accelerate a given payload to a given delta-v, the methane stage will be more compact and have a lower dry mass.

If we assume ambitious numbers for both hydrolox (470s isp, 90% prop. mass fraction) and methalox (380s isp, 93% pmf), hydrolox stages have a lower dry mass from 5.6km/s upwards.

A point of trivia the Vulcan ACES PF is 96% due to the extremely light weight stainless steel balloon  tanks. You could get even higher PF's for MethaLox.

In addition its not the necessarily the deltaV you can get from the systems but the energy you can get from the prop as a payload that is delivered from somewhere to somewhere else. Its this that makes the operational costs of an in-space hydrolox system have nearly half the prop costs than that of a MethaLox in-space system if it cost the same per kg for both propellants, which it doesn't MethaLox would require significantly more energy for its mining/production except on Mars. Meaning use of MethaLox is only good cost wise until hydrolox is available in-space in quantity for cheaper than it costs to transport it from Earth or from Mars surface into space.

In a highly reused in-space system almost all costs are the purchase of the prop 150mt of hydroLox for a Vulcan ACES to move 100mt of payload at $1000/kg is $150M and to do the same 100mt of payload movement with a Raptor 220mt of prop at the same $1000/kg costs $220M. Its the costs of the prop not the costs of the stages or their efficiency or anything else.

[Robotbeat]

To accelerate a given payload to a given delta-v, the methane stage will be more compact and have a lower dry mass.

If we assume ambitious numbers for both hydrolox (470s isp, 90% prop. mass fraction) and methalox (380s isp, 93% pmf), hydrolox stages have a lower dry mass from 5.6km/s upwards.

More like 95% for methalox if hydrolox is 90%. You get higher T/W ratio for the engine due to using hydrocarbon and you can fit twice as much propellant in, so it comes out close to being twice as good for mass ratio. ...though of course, you have to include payload mass. But on the way back, if you're reusing the stage (say, in-space ACES-style), then you're only pushing the stage's dry mass around. This is partly why dry mass is especially important for reusability (versus Isp) than it would be for expendable.

[Oli]

A point of trivia the Vulcan ACES PF is 96%

No way.

In a highly reused in-space system almost all costs are the purchase of the prop

If you deliver propellant to staging points with SEP, ISP becomes less and storability more important. But I suppose you're correct in principle.

More like 95% for methalox if hydrolox is 90%.

Is that speculation or do you have a source? I got my numbers from the Soyuz 5 thread.

You get higher T/W ratio for the engine due to using hydrocarbon and you can fit twice as much propellant in, so it comes out close to being twice as good for mass ratio.

It's not that simple.

[Robotbeat]

...

More like 95% for methalox if hydrolox is 90%.

Is that speculation or do you have a source? I got my numbers from the Soyuz 5 thread.

That's not a better source.

You get higher T/W ratio for the engine due to using hydrocarbon and you can fit twice as much propellant in, so it comes out close to being twice as good for mass ratio.

It's not that simple.

Sure it is. The bulk density is actually MORE than twice as good for methane/LOx, at 828kg/m^3 for methane/LOx and 358kg/m^3 for hydrolox: http://web.archive.org/web/20130515142359/http://www.dunnspace.com/alternate_ssto_propellants.htm
...which is a factor of 2.3. Also, you need more insulation for hydrogen, and you'll still have faster boil-off... Both those things add mass for hydrolox. Things like pressurization system mass are roughly proportional to volume (due to the pressure vessel equation). So yeah, there are more details, but you're quite close if you just use a factor of 2.

Density matters. A lot.^TM

[Oli]

That's not a better source.

A better source than what? You have not even given one.

but you're quite close if you just use a factor of 2.

"quite close" is not good enough.