Advanced propulsive fluid accumulator concepts and propellant combinations

[zypofaeser]

There have previously been some discussions about the possibility of using propulsive fluid accumulation in orbit, and the advantages that this might have compared to nuclear thermal rockets (More specifically in the thread about the DRACO rockets). Therefore I thought it would be useful to have a thread dedicated to such concepts.

Prior concepts involved the accumulation of liquid oxygen in low earth orbit, on a satellite which would later act as a propellant depot, where an upper stage could resupply with liquid oxygen. For an Earth Departure Stage, you could significantly reduce the mass to be launched to LEO, if you were going to launch toward a higher orbit, Luna, Mars, etc.

However, this has the obvious issue that a decent fraction of your propellant would still have to be carried from Earth or elsewhere. As an example the RL10 is often listed as having an oxidizer to fuel ratio of 5.88 to 1, meaning that 14,5% of your propellant would still come from Earth. This is low, but as you will see, it is far from the lowest we can go.

The main trick that we can use is exploiting the nitrogen in the atmosphere to "bulk up" the propellant with only the hydrogen coming from Earth. If we look at flight proven fuel mixtures such as ammonia/LOx, which was used on the X15, you will see some improvement to around 7,3%, at the cost of a significant reduction in your specific impulse. However, I believe that we can go further. Using nitrous oxide instead of liquid oxygen would provide even more "bulk" to the propellant mixture, even if it doesn't bring much more energy to the table. This doesn't matter much, as the great improvement in the possible mass ratio should overcome many of the issues of a reduced specific impulse. The propellant could, at least theoretically, need as little as 3,6% of its mass to be delivered.

As an example, assuming equal dry masses, a rocket doing a 4500m/s manoeuvre would need a mass ratio of 2,78 (assuming an exhaust velocity of 4400m/s), which means that for every ton of rocket you would need 0,26 tons of hydrogen from Earth. With the LOx/ammonia rocket (assuming an exhaust velocity of 3368 m/s as listed on the Encyclopedia Astronautica), your mass ratio would be 3,80 with only 0,205 being delivered, but with N2O/ammonia you would get a mass ratio of 7,07, but with only 0,219 being delivered, assuming an exhaust velocity of only 2300m/s.

On lower delta-v missions, you would be better of using the N2O/ammonia rockets, whereas on higher delta-v missions, you would benefit from having an upper stage using hydrolox engines. With the N2O/ammonia rocket, you also have the advantage that the propellants are not hard cryogens. While both would require some cooling, it is a much more achievable temperature region, especially if stored under pressure, such as in a pressure fed rocket engine. This would be useful if you're filling up a depot, while waiting for a launch window.

An example of an application of this technology would be as a kick stage to an interplanetary craft. A Starship, assumed to be able to carry 70 tons of LH2 per launch (constrained by volume), would be able to produce 1944 tons of N2O/ammonia. On a 4500m/s mission as listed above (a trans Mars injection), you could propel a spacecraft weighing 320 tons (including the rocket). This makes it fairly reasonable to think that a kick stage like this could be used to reduce the number of launches needed for a given payload mass on Mars, whether this is by Starship, or another mission architecture. Whether it will be economically viable or not, especially when considering the cost of acquiring the propellant in LEO, is an open question.

Sources, just sauce, no ketchup:
https://en.wikipedia.org/wiki/Propulsive_fluid_accumulator
https://en.wikipedia.org/wiki/RL10#Table_of_versions
http://www.astronautix.com/l/loxammonia.html

[LMT]

There have previously been some discussions about the possibility of using propulsive fluid accumulation in orbit, and the advantages that this might have...

Whether it will be economically viable or not... is an open question.

I gave some info to address open questions, when you asked.  You skipped over the info.  Odd.

[zypofaeser]

I didn't skip the info provided, it just simply didn't answer all relevant questions, thus requiring further investigations and discussion.

I would agree that hydrazine is undesirable, which is why post mentions ammonia as a propellant instead, mainly because ammonia is less toxic (although far from being completely safe), and the fact that ammonia is more stable. Otherwise you might need to add in MMH or UDMH to keep the hydrazine usable for regenerative cooling.

However, while I believe that we both agree that there is a significant benefit to be had if this tech works, my main concern is not with the toxicity and the handling, but with the manufacturing in orbit. How much will a given PROFAC unit be able to acquire in a given time, and how long will it last? If the answer is that a PROFAC unit will yield several tons of propellant per ton of PROFAC unit, then it will likely be succesful, no matter the specific fuel used. However, if it only produces say 0,25 tons per ton every year and it has a lifespan of only 10 years, you're probably better off just investing in another reusable rocket to launch methalox directly, given the cost of the PROFAC unit and the long time to acquire the propellant.

Also, I chose to base my calculations on ammonia/nitrous oxide because they're stable without solvents. HAN (hydroxylammonium nitrate) requires a solvent, as pure solid HAN is explosive and AFAIK it would result in disaster if used as a fuel in its pure form. For further details see https://haz-map.com/Agents/14825

[LMT]

I didn't skip the info provided, it just simply didn't answer all relevant questions, thus requiring further investigations and discussion.

I would agree that hydrazine is undesirable, which is why post mentions ammonia as a propellant instead, mainly because ammonia is less toxic (although far from being completely safe), and the fact that ammonia is more stable. Otherwise you might need to add in MMH or UDMH to keep the hydrazine usable for regenerative cooling.

However, while I believe that we both agree that there is a significant benefit to be had if this tech works, my main concern is not with the toxicity and the handling, but with the manufacturing in orbit. How much will a given PROFAC unit be able to acquire in a given time, and how long will it last? If the answer is that a PROFAC unit will yield several tons of propellant per ton of PROFAC unit, then it will likely be succesful, no matter the specific fuel used. However, if it only produces say 0,25 tons per ton every year and it has a lifespan of only 10 years, you're probably better off just investing in another reusable rocket to launch methalox directly, given the cost of the PROFAC unit and the long time to acquire the propellant.

Also, I chose to base my calculations on ammonia/nitrous oxide because they're stable without solvents. HAN (hydroxylammonium nitrate) requires a solvent, as pure solid HAN is explosive and AFAIK it would result in disaster if used as a fuel in its pure form. For further details see https://haz-map.com/Agents/14825

A lot of mistakes there. 

HAN was formulated as AF-M315E monopropellant and is marketed as ASCENT.  It's neither a "solid" nor an aqueous solution but, again, an ionic liquid (liquid salt).  You remove water to make ionic liquids. 

"Explosive", yes, like other propellants.  5 years in space missions now.

[edzieba]

"Explosive", yes, like other propellants.

False. There are plenty of propellants that are non-explosive when stored, including the majority of non-hypergolic bipropellants. Trivial examples include hybrids (the solid fuel grain is nonexplosive), Kerolox (you need to do quite a bit of work mixing RP-1 with an oxidiser to get to to explode), etc. This is not the case for monopropellants (e.g. HAN) which do not require mixing with an oxidisers to be explosive.

In addition, AF-M315E is not water free. It is 11% water as standard, and can still ignite up to at least 60% water content (AF-M315EM) - image from "Non-Catalytic Microwave Ignition of Green Hydrazine Replacements".

[Vultur]

I think the big question with this kind of technology is whether the cost is worth it if Starship-type orbital propellant transfer is developed.

I used to be a big fan of the idea, but if Starship tankers can launch propellant for like $100/kg or less... you'd have to get a *lot* of propellant out of the accumulator over its lifetime to make it worth its development and build costs.

[LMT]

AF-M315E is not water free.  It is 11% water as standard...

That's an F-16 EPU test, not a space reference.  They diluted to "reduce its combustion temperature," on old turbine blades.

Cf. GR-1A thruster.

[LMT]

I think the big question with this kind of technology is whether the cost is worth it if Starship-type orbital propellant transfer is developed.

I used to be a big fan of the idea, but if Starship tankers can launch propellant for like $100/kg or less... you'd have to get a *lot* of propellant out of the accumulator over its lifetime to make it worth its development and build costs.

To start, a drag-compensated Starship trawler could conceivably collect 860 tons of LOX over 3 weeks, at 90 km altitude.  At fleet scale, savings add up.

[Vultur]

I think the big question with this kind of technology is whether the cost is worth it if Starship-type orbital propellant transfer is developed.

I used to be a big fan of the idea, but if Starship tankers can launch propellant for like $100/kg or less... you'd have to get a *lot* of propellant out of the accumulator over its lifetime to make it worth its development and build costs.

To start, a drag-compensated Starship trawler could conceivably collect 860 tons of LOX over 3 weeks, at 90 km altitude.  At fleet scale, savings add up.

It's only savings if the cost of collecting propellant this way is less than the cost of launching it on tankers.

 What do you think the marginal cost of launching a tanker will be once they're reusable? If it's like $5 million then the cost of propellant in LEO would be something like $33/kg (assuming 150,000 kg per tanker). The payback time for developing any new way to get propellant would be really long.

[LMT]

you'd have to get a *lot* of propellant out of the accumulator over its lifetime to make it worth its development and build costs.

To start, a drag-compensated Starship trawler could conceivably collect 860 tons of LOX over 3 weeks, at 90 km altitude.  At fleet scale, savings add up.

It's only savings if the cost of collecting propellant this way is less than the cost of launching it on tankers.

What do you think the marginal cost of launching a tanker will be once they're reusable? If it's like $5 million then the cost of propellant in LEO would be something like $33/kg (assuming 150,000 kg per tanker). The payback time for developing any new way to get propellant would be really long.

Not long as shown, no.  Suitable technologies, such as 3M Nextel ceramic textiles and the SABRE cooler/compressor, exist. 

At your notional pricing:

-  One baseline LOX trawler would equate with half a $ billion per year. 

-  Or more, if you trawled LN2 concurrently. 

-  And when you consider the vanished infrastructure costs, you see even greater benefits.

It's a timely topic, especially for those with interest in the terrible economics of interplanetary fleets.

[Vultur]

you'd have to get a *lot* of propellant out of the accumulator over its lifetime to make it worth its development and build costs.

To start, a drag-compensated Starship trawler could conceivably collect 860 tons of LOX over 3 weeks, at 90 km altitude.  At fleet scale, savings add up.

It's only savings if the cost of collecting propellant this way is less than the cost of launching it on tankers.

What do you think the marginal cost of launching a tanker will be once they're reusable? If it's like $5 million then the cost of propellant in LEO would be something like $33/kg (assuming 150,000 kg per tanker). The payback time for developing any new way to get propellant would be really long.

Not long as shown, no.  Suitable technologies, such as 3M Nextel ceramic textiles and the SABRE cooler/compressor, exist. 

The tech may exist, but making a Starship based trawler still requires more development than launching more tankers of an existing mass produced design. If propellant in LEO is this cheap, it's hard to justify any propellant based development.

At your notional pricing:

-  One baseline LOX trawler would equate with half a $ billion per year. 

Meaning it would produce over 15 million tons of LOX per year? That seems shockingly high.

-  Or more, if you trawled LN2 concurrently. 

What's the use of LN2 as propellant in a methalox based architecture?

-  And when you consider the vanished infrastructure costs, you see even greater benefits.

I am unconvinced building infrastructure for more Starship launches (assuming Starship is already up and running) is terribly expensive.

It's a timely topic, especially for those with interest in the terrible economics of interplanetary fleets.

Why are they terrible? (The traditional argument is that the long synod cycle means few reuses, but if Starship is mass-produced cheaply, that isn't necessarily a terrible thing. I think that mindset requires spacecraft to be near-unique, boutique production models.  Most Starships going to Mars will probably never leave it, never be reused... but that is ok if they are cheap to make.)

[LMT]

At your notional pricing:

-  One baseline LOX trawler would equate with half a $ billion per year. 

Meaning it would produce over 15 million tons of LOX per year? That seems shockingly high.

$33/kg

[Vultur]

At your notional pricing:

-  One baseline LOX trawler would equate with half a $ billion per year. 

Meaning it would produce over 15 million tons of LOX per year? That seems shockingly high.

$33/kg

Oh, sorry. 15,000 tons per year. Oops.

[edzieba]

AF-M315E is not water free.  It is 11% water as standard...

That's an F-16 EPU test, not a space reference.  They diluted to "reduce its combustion temperature," on old turbine blades.

Cf. GR-1A thruster.

Did you read your own citation?

[LMT]

AF-M315E is not water free.  It is 11% water as standard...

That's an F-16 EPU test, not a space reference.  They diluted to "reduce its combustion temperature," on old turbine blades.

Cf. GR-1A thruster.

Did you read your own citation?

Thrusters should be understood in a spaceflight forum.  Again, cf. the new GR-1A thruster.  What happens to Isp when you remove water?

Due to enhanced heat transfer in small scale system, the new GR-M1 thrusters run on a 10 wt% water diluted AF-M315E propellant, which reduces the specific impulse to 206 s (as compared to 231 s in GR-1A thruster).

[edzieba]

The GR-1A also uses the same propellant mix - the difference in ISP between the GR-1A and the GR-M1 is due to the size of the thruster (the GR-1A being twice the size of the GR-M1). See attached image 1.

There is no 'dry' AF-M315E. It's 10% H20 as standard, with further dilution being an option. There are also no other 'dry' HAN mixes. See attached image 2.

This is not a surprise, as reaction of HAN with water is the first step in all the reaction chains of HAN monopropellants. See attached image 3.

[LMT]

the difference in ISP between the GR-1A and the GR-M1 is due to the size of the thruster (the GR-1A being twice the size of the GR-M1).

No, they explained the Isp difference, right above.  Dilution is the straightforward explanation.  Don't ignore info.

reaction of HAN with water is the first step in all the reaction chains of HAN monopropellants.

If you think water is needed for an ammonium nitrate explosion, you probably shouldn't transport fertilizer.

Further confirmation of my previous posts:

To partially mitigate thermal management challenges exacerbated at the miniature scale, the GR-M1 is designed to operate on a reduced-flame-temperature variant of the ASCENT propellant containing 10% added water.

The OP was propellant accumulation in orbit?

Refs.

Yost, B. and Weston, S., 2024. State-of-the-art small spacecraft technology (No. NASA/TP-20240001462).

[edzieba]

the difference in ISP between the GR-1A and the GR-M1 is due to the size of the thruster (the GR-1A being twice the size of the GR-M1).

No, they explained the Isp difference, right above.  Dilution is the straightforward explanation.  Don't ignore info.

The two thrusters use the exact same propellant, as shown in the table attached to the prior post. They do not use any different 'dilution', they use the exact same propellant mix, which contains 10% water as a necessity to work. As mentioned above, presence of water is a requirement of HAN-based propellants. No water, no reaction, no propellant.
From https://doi.org/10.1016/j.actaastro.2022.04.011

HAN reaction is suppressed by presence of excess water or absence of water

If you think water is needed for an ammonium nitrate explosion, you probably shouldn't transport fertilizer.

HAN is not Ammonium Nitrate. HAN is hydroxylamine nitrate. Two different chemicals.

In addition, the desired behaviour of a propellant is controlled stable combustion, not unstable explosive decomposition.

[LMT]

the difference in ISP between the GR-1A and the GR-M1 is due to the size of the thruster (the GR-1A being twice the size of the GR-M1).

No, they explained the Isp difference, right above.  Dilution is the straightforward explanation.  Don't ignore info.

The two thrusters use the exact same propellant, as shown in the table attached to the prior post. They do not use any different 'dilution', they use the exact same propellant mix, which contains 10% water as a necessity to work.

No, the old GR-M1 dilution is plainly stated, and understood. 

Don't clutter / derail, but focus on the OP.

[edzieba]

the difference in ISP between the GR-1A and the GR-M1 is due to the size of the thruster (the GR-1A being twice the size of the GR-M1).

No, they explained the Isp difference, right above.  Dilution is the straightforward explanation.  Don't ignore info.

The two thrusters use the exact same propellant, as shown in the table attached to the prior post. They do not use any different 'dilution', they use the exact same propellant mix, which contains 10% water as a necessity to work.

No, the old GR-M1 dilution is plainly stated, and understood. 

Same water percentage by weight, see attached (again). Both use AF-315E, and AF-315E has 10% water. That's a fundamental part of the propellant mix without which it will not function - no difference in 'dilution', as this is not dilution but a core component of HAN monoprop mixes that cannot be omitted. More water can be added if desired to tune the mix for other applications, but AF-315E used in both the GR-1A and GR-M1 thrusters has 10% water.

We can curtail the idea of a waterless HAN propellant, as such a thing does not and cannot exist.