Should launch providers switch to common methalox engines?

[Pipcard]

Musk says that overhead starts with how the launch vehicle is designed. The workhorse Atlas V, for example, used for everything from planetary probes to spy satellites, employs up to three kinds of rockets, each tailored to a specific phase of flight. The Russian-built RD-180 first- stage engines burn a highly refined form of kerosene called RP1. Optional solid-fuel strap-on boosters can provide additional thrust at liftoff, and a liquid hydrogen upper stage takes over in the final phase of flight. Using three kinds of rockets in the same vehicle may optimize its performance, but at a price: “To a first-order approximation, you’ve just tripled your factory costs and all your operational costs,” says Musk.

Instead, from the very beginning, SpaceX designed its Falcon rockets with commonality in mind. Both of Falcon 9’s stages are powered by RP1 and liquid oxygen, so only one type of engine is required. Both are the same diameter and are constructed from the same aluminum-lithium alloy, reducing the amount of tooling and the number of processes and resulting in what Musk calls “huge cost savings.”

If that is true, if having three different types of engines really triples the operating costs...

then why aren't others thinking of "optimizing for (manufacturing and operating) costs" for their next-generation launch vehicles by switching to a common propellant?

Let's say there was a launch provider that used a hydrolox+kerolox+solid combination. They managed to become a major player in the launch industry (case in point: Ariane 4 used multiple types of engines and still took over 50% of the market). Now let's say they were designing a next-generation rocket, and that they realized that they should optimize for cost through commonality, in order to stay competitive.

If they were to switch to an all-hydrolox rocket (without any solids), the first stage wouldn't be optimal because of liquid hydrogen's low density and the relatively low thrust of hydrolox engines. It is also apparently "hard to handle," which was why SpaceX abandoned the idea.

If they were to switch to an all-kerolox rocket, they would sacrifice the performance given by the high-energy upper stage.

Methalox is not the most optimal for either stage. But it is a compromise. Its density and specific impulse are in-between that of kerolox and hydrolox. There hasn't been a methalox engine used in any rocket before, but judging by the in-development Raptor, it can have a good thrust for first stages, too. SpaceX is planning for their next generation launcher (BFR) to be all-methalox. It has the additional benefit with being common with Mars ISRU and having less residue buildup than kerolox, good for reusability. The reason that they didn't start out with it is that it was easier for them, as a startup, to develop kerolox engines.

For major launch providers based in countries that, in general, have relatively higher labor costs (i.e. not China, Russia, or India):
- Ariane 6 and H-III will have hydrolox and solids because of heritage (less development costs).
- Vulcan will have methalox (because it's an "in-between" fuel for the first stage), hydrolox, and solids (the latter two also because of heritage).

The main thing I want to ask is, is it worth it for a provider to abandon their hydrolox/kerolox/solid manufacturing and processing infrastructure and replace that with a methalox-based one? I know that solid rocket manufacturing can have synergy with the defense industry (i.e. missiles), though.

[woods170]

Musk says that overhead starts with how the launch vehicle is designed. The workhorse Atlas V, for example, used for everything from planetary probes to spy satellites, employs up to three kinds of rockets, each tailored to a specific phase of flight. The Russian-built RD-180 first- stage engines burn a highly refined form of kerosene called RP1. Optional solid-fuel strap-on boosters can provide additional thrust at liftoff, and a liquid hydrogen upper stage takes over in the final phase of flight. Using three kinds of rockets in the same vehicle may optimize its performance, but at a price: “To a first-order approximation, you’ve just tripled your factory costs and all your operational costs,” says Musk.

Instead, from the very beginning, SpaceX designed its Falcon rockets with commonality in mind. Both of Falcon 9’s stages are powered by RP1 and liquid oxygen, so only one type of engine is required. Both are the same diameter and are constructed from the same aluminum-lithium alloy, reducing the amount of tooling and the number of processes and resulting in what Musk calls “huge cost savings.”

If that is true, if having three different types of engines really triples the operating costs...

then why aren't others thinking of "optimizing for (manufacturing and operating) costs" for their next-generation launch vehicles by switching to a common propellant?

The answer is simple. Until just a few years ago nobody took SpaceX seriously. There was no need for "optimizing for cost". Now that SpaceX has begun taking bites from the revenues of the usual suspects, the need for "optimizing for cost" is slowly but ever so gradually beginning to sink in with the rest of the industry.

[Rocket Science]

My problem with this is if an issue with the engine crops up all the launchers using it would be grounded...

[nadreck]

The main thing I want to ask is, is it worth it for a provider to abandon their hydrolox/kerolox/solid manufacturing and processing infrastructure and replace that with a methalox-based one? I know that solid rocket manufacturing can have synergy with the defense industry (i.e. missiles), though.

This depends on the volume of business you will be doing, the amount of extra capital needed to switch, and the difference in unit cost at the end. If the company currently launches 10 of the hydrolox/kerolox/solid combination vehicles and owns the plants that make the engines and tanks (not the actual case in the real world) then one might assume that there is less retooling expense when they go to a new design that uses the same mix of propulsion than replacing everything with methalox. So if the new vehicle would cost $100M each and require an extra $1B in capital in retooling if it was a methalox rocket, but instead cost $120M each if it stayed hydrolox/kerolox/solid, then it would, at 10 flights a year, take 5 years to pay off the difference in up front capital at a 0% discount rate and about 6 years at a 5% discount rate. If the difference was either $2B in capital or the hydrolox/kerolox/solid rocket only cost $110M then it would take 10 years at a 0% discount rate, or more than 15 years at a 5% discount rate. It really might not be worth going to methalox if it was 15 years before the benefits would start to pay off as you might be into a redesign by then. Double the flight rate and it all changes back to the first case, halve  the flight rate and the first case looks like the 2nd.

[Oli]

Musk says that overhead starts with how the launch vehicle is designed. The workhorse Atlas V, for example, used for everything from planetary probes to spy satellites, employs up to three kinds of rockets, each tailored to a specific phase of flight. The Russian-built RD-180 first- stage engines burn a highly refined form of kerosene called RP1. Optional solid-fuel strap-on boosters can provide additional thrust at liftoff, and a liquid hydrogen upper stage takes over in the final phase of flight. Using three kinds of rockets in the same vehicle may optimize its performance, but at a price: “To a first-order approximation, you’ve just tripled your factory costs and all your operational costs,” says Musk.

Instead, from the very beginning, SpaceX designed its Falcon rockets with commonality in mind. Both of Falcon 9’s stages are powered by RP1 and liquid oxygen, so only one type of engine is required. Both are the same diameter and are constructed from the same aluminum-lithium alloy, reducing the amount of tooling and the number of processes and resulting in what Musk calls “huge cost savings.”

If that is true, if having three different types of engines really triples the operating costs...

We don't have enough data to say whether that is true or not. From a development cost point of view he's certainly right though. There's no way SpaceX could have afforded to develop RD-180, RL-10, AJ-62 equivalents.

[nadreck]

Musk says that overhead starts with how the launch vehicle is designed. The workhorse Atlas V, for example, used for everything from planetary probes to spy satellites, employs up to three kinds of rockets, each tailored to a specific phase of flight. The Russian-built RD-180 first- stage engines burn a highly refined form of kerosene called RP1. Optional solid-fuel strap-on boosters can provide additional thrust at liftoff, and a liquid hydrogen upper stage takes over in the final phase of flight. Using three kinds of rockets in the same vehicle may optimize its performance, but at a price: “To a first-order approximation, you’ve just tripled your factory costs and all your operational costs,” says Musk.

Instead, from the very beginning, SpaceX designed its Falcon rockets with commonality in mind. Both of Falcon 9’s stages are powered by RP1 and liquid oxygen, so only one type of engine is required. Both are the same diameter and are constructed from the same aluminum-lithium alloy, reducing the amount of tooling and the number of processes and resulting in what Musk calls “huge cost savings.”

If that is true, if having three different types of engines really triples the operating costs...

Exactly, if you are the incumbent with sunk costs (but really no except SpaceX fully is since they all buy the different engines from an array of other companies) then it makes less sense to change unless you are sure of the volume, but if you are the new guy then whatever you choose you have to tool up for so you have incentive to minimize the manufacturing capitalization.

We don't have enough data to say whether that is true or not. From a development cost point of view he's certainly right though. There's no way SpaceX could have afforded to develop RD-180, RL-10, AJ-62 equivalents.

[muomega0]

The answer depends on the launch manifest and destination and long term plans.  Two different engines exist to avoid common failure modes.

ISP rules in space, so depending on the deltaV one can increase IMLEO by 25% using methane vs LH2, so is the NASA 'market' 0, 50, or 500mT/year?  The relative importance of I_sp and propellant density depends on the application (e.g., first stage or upper stage) and the mission (e.g., heavy payload to low orbit or small payload to escape).

Another consideration is the common configuration for Class A cargo and crew to find that unknown unknown.  Perhaps solids are not used for crew due to LAS mass, but are used for performance gains in the heavy variant. 

Another consideration is methane vs water for ISRU, and choosing the former pretty much eliminates LH2 option.  Type of fuel for landers, if any?

If launching mass to BEO is not in the future plans, then its entirely driven by DOD needs, which has what, 8 heavy flights over 13 years, and lacks cash for space assets.   

In this case, the choices are narrowed further to how many $ are available in the short term, and even further to those with semi-deep pockets, who can fund, err, 'commercial' engines and LVs.

Another consideration is where the components are built, where they are launched, and how many $$ can be diverted from state funding to tax breaks, like using rockets to round up cows or planting farms around pads.

Long term planning for NASA and the USG is not possible until campaign finance and gerrymandered districts are corrected.   It will be interesting to see what answer industry tells the USG, but right now its Vulcan v0 and Falcon.

[Jim]

The main thing I want to ask is, is it worth it for a provider to abandon their hydrolox/kerolox/solid manufacturing and processing infrastructure and replace that with a methalox-based one? I know that solid rocket manufacturing can have synergy with the defense industry (i.e. missiles), though.

No.  Do all land vehicles use the same fuel?

[rayleighscatter]

I'm not sure I'd want to put all my eggs in the methane basket, it's becoming a real environmental bugaboo.

[nadreck]

I'm not sure I'd want to put all my eggs in the methane basket, it's becoming a real environmental bugaboo.

I'm not sure I'd want to put all my eggs in the methane basket, it's becoming a real environmental bugaboo.

It has a much smaller GHG footprint than any other fossil fuel between how it is gathered and burned. The fact that it in and of itself is a greenhouse gas is irrelevant as the vast majority of it that is removed from the ground (more than 99.9%) is burned to become mainly water, half as much CO₂, and in the case of other applications there are trace amounts of other chemicals being burned as well as nitrogen combining in as it is being burned giving us some other pollutants, but when you use it in a rocket engine it is presumably more refined with just traces of ethane left in and no nitrogen is being mixed in with it during combustion.

[rayleighscatter]

I'm not sure I'd want to put all my eggs in the methane basket, it's becoming a real environmental bugaboo.

I'm not sure I'd want to put all my eggs in the methane basket, it's becoming a real environmental bugaboo.

It has a much smaller GHG footprint than any other fossil fuel between how it is gathered and burned. The fact that it in and of itself is a greenhouse gas is irrelevant as the vast majority of it that is removed from the ground (more than 99.9%) is burned to become mainly water, half as much CO₂, and in the case of other applications there are trace amounts of other chemicals being burned as well as nitrogen combining in as it is being burned giving us some other pollutants, but when you use it in a rocket engine it is presumably more refined with just traces of ethane left in and no nitrogen is being mixed in with it during combustion.

Don't underestimate the US government, especially when methane has already been singled out for stricter scrutiny from production to transportation to storage and use. It might seem absurd, but few also thought in 1960 that barely a decade later their brand new jet aircraft would be banned for being too noisy.

[Nilof]

I'm not sure I'd want to put all my eggs in the methane basket, it's becoming a real environmental bugaboo.

I'm not sure I'd want to put all my eggs in the methane basket, it's becoming a real environmental bugaboo.

It has a much smaller GHG footprint than any other fossil fuel between how it is gathered and burned. The fact that it in and of itself is a greenhouse gas is irrelevant as the vast majority of it that is removed from the ground (more than 99.9%) is burned to become mainly water, half as much CO₂, and in the case of other applications there are trace amounts of other chemicals being burned as well as nitrogen combining in as it is being burned giving us some other pollutants, but when you use it in a rocket engine it is presumably more refined with just traces of ethane left in and no nitrogen is being mixed in with it during combustion.

Keep in mind though, that to get optimal Isp, methalox has to burn fuel-rich. So you are releasing the excess methane directly into the atmosphere.

The rocket company (the book) had a methane RLV and this issue did turn up once reuse became common, with the trades involved when switching to an oxidizer-rich mode.

[Steven Pietrobon]

Keep in mind though, that to get optimal Isp, methalox has to burn fuel-rich. So you are releasing the excess methane directly into the atmosphere.

The oxygen in the atmosphere will quickly burn with the very hot excess methane. The main problem is boil off on the ground. That is solved by piping the gaseous methane to a burn off point, just like what is done for hydrogen boil off.

As to the question of the thread, the rocket builder should be free to use whatever propellant combination they think is best, although moving away from toxic propellants should be encouraged.

[Proponent]

Don't underestimate the US government, especially when methane has already been singled out for stricter scrutiny from production to transportation to storage and use. It might seem absurd, but few also thought in 1960 that barely a decade later their brand new jet aircraft would be banned for being too noisy.

Given that unburnt methane is a powerful greenhouse gas, I don't see what's fundamentally irrational about paying close attention to the parts of the methane cycle leading up to combustion.  Likewise, since commercial aviation grew rapidly in the decade after 1960, it's not obvious to me that jet-engine noise might not have gone from being a minor annoyance to a significant problem over the same period.

[Robotbeat]

Hydrogen is also an indirect greenhouse gas that by mass is worse than CO2 (not by molar mass or volume, though, if I recall correctly). But like methane, that's only if it's unburnt.

This is all overblown. There are inexpensive ways to counteract the effect, such as burying biochar. It's not significant.

[Lars-J]

To answer the question posed by the thread starter: NO.

Think of it this way... Should early car manufacturers have been forced to use one or two shared engines? Of course not. That would not have benefitted anyone.

[Pipcard]

To answer the question posed by the thread starter: NO.

Think of it this way... Should early car manufacturers have been forced to use one or two shared engines? Of course not. That would not have benefitted anyone.

It's not about sharing the same engine between launch providers, it's about each provider having a common engine & propellant for their rocket stages to "optimize for cost."

(I'm sorry if you read the title wrong)

[Stan-1967]

Keep in mind though, that to get optimal Isp, methalox has to burn fuel-rich. So you are releasing the excess methane directly into the atmosphere.

The rocket company (the book) had a methane RLV and this issue did turn up once reuse became common, with the trades involved when switching to an oxidizer-rich mode.

How do you arrive at the conclusion that methane, in any significant amount, will being released into the air?  Any hydrocarbon undergoing combustion can have intermediate carbon compounds like CO as they combust fully to CO2.   It is not supportable that a fuel rich cycle will somehow miraculously fully combust some portion of CH4 completely to C02 and H20 and the balance of the CH4 will not experience any combustion at all?

The combustion gasses will be primarily a mixiture of CO2, H20, & CO, with CH4 being the least plausible of all emitted gasses. CH3, 0, OH, & CH20 are all more probable than CH4. 

[TrevorMonty]

Most Hydrogen is extracted from methane with CO2 being by product.
Then there is shipping and extra energy required to cool LH. As a rocket fuel is methane is more environmentally friendly if burnt direct compared to converting to LH.
 
https://en.m.wikipedia.org/wiki/Hydrogen_production

[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.

[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.

[Robotbeat]

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.

Ive explained my reasoning in sufficient detail and provided a link to supporting analysis (if you want, I can point to the pressure vessel equation, but it just says vessel dry mass is proportional to volume). You haven't supplied a link or explained reasoning in any detail. The ball is in your court.

[TrevorMonty]

A point of trivia the Vulcan ACES PF is 96%

No way.

 Centuar is 90% with IVF expected to increase that to low 90s. So 96% for ACES may not be far off the mark.

[Oli]

Ive explained my reasoning in sufficient detail and provided a link to supporting analysis (if you want, I can point to the pressure vessel equation, but it just says vessel dry mass is proportional to volume). You haven't supplied a link or explained reasoning in any detail. The ball is in your court.

First, my numbers come from this post, which is a methalox rocket design from actual Russian engineers. The upper stage has a pmf of 92%, but I was nice and picked the pmf of the first stage which is 93% (both rounded).

Your "source" does a rather unsophisticated calculation of payload for a constant-volume SSTO. They say the model should be used to compare fuels with roughly similar density/isp. More useful formulas for dry mass calculation can be found here. In-space stages are less volume-limited than SSTO (no drag, TPS).

Centuar is 90% with IVF expected to increase that to low 90s. So 96% for ACES may not be far off the mark.

In terms of actual mass savings its a huge step.

Centaur has the best pmf of any hydrolox upper stage to date. Of course there's always better stuff out there, e.g. the Ariane 5 core, but I tried to make a fair comparison.
 
If you look at NASA designs for in-space stages, they usually have far worse pmfs than upper stages.

[Robotbeat]

Ive explained my reasoning in sufficient detail and provided a link to supporting analysis (if you want, I can point to the pressure vessel equation, but it just says vessel dry mass is proportional to volume). You haven't supplied a link or explained reasoning in any detail. The ball is in your court.

First, my numbers come from this post, which is a methalox rocket design from actual Russian engineers. The upper stage has a pmf of 92%, but I was nice and picked the pmf of the first stage which is 93% (both rounded).
...

Russians generally build their rocket stages like tanks. You can see people walk on top of Soyuz boosters as they're being built. In contrast, the American hydrolox stages like Centaur that can achieve 90% PMF are built with balloon tanks, with metal so thin it can't even support its own weight under gravity and must be pressurized at all times. You're using a false comparison.

(This isn't a dig at Russian engineering. Building their rockets like that has served them well.)

[Pipcard]

What they [Arianespace] don't have is the culture of a SpaceX, which allows them to change the ways to adapt quickly with what might be heretical approaches that challenge technology base and heritage - because evolving a hydrolox propulsion system to a less expensive, modular launch architecture (possibly reusable) requires addressing massive changes of esoteric nature with a KISS approach that can be replicated with lowest labor costs, in an environment that is motived by entirely the opposite mindset.

Kerolox and methalox launchers with a single engine type make for the most economical LV - this simply won't go away. Forget for the moment the reusability aspect - just from the standpoint of supporting reliable production through the smallest footprint, approaches like this win at the budget level, but compromise at the launch vehicle performance level. If you can't accept the performance compromise(including flying multiple launch missions) then one must accept the burden of 10x budget (or more) for what it takes for optimal propulsion.

If a launch vehicle manufacturer had already made investments into developing "high-energy" hydrolox stages, should they abandon that, or should they retain that knowledge and capability?

[jongoff]

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 actually may not be correct, based on a conversation I had with Paul Spudis at a conference this last week that I summarized in this post:

http://forum.nasaspaceflight.com/index.php?topic=39559.msg1612453#msg1612453

tl;dr version is that he says of the two sensors that measured concentrations, the one that they have much more confidence in was saying 90-95% water. The sensor that was revised down to show less water was one that we don't have a lot of confidence in our ability to interpret.

~Jon