amount of gas needed to keep up mass flow rate

[Thunder Chicken]

Ok, from what I gather, I'd have to multiply m^3/s * Pa, so that would be equivalent to litres/s * kPa. Which leaves me with 41 kW, before factoring in efficiency. That's a small car! No idea how I'd power that from batteries...

41 kW really isn't terrible. The Saturn V first stage fuel pump was something like 40 MW. A Tesla car battery can receive or transfer power at rates from 450-650 kW.

The trick is that you need to determine the amount of energy that needs to be transferred - how long you need to maintain this power. Take your 41 kW, divide by pump efficiency, and multiply it by the number of hours you need to operate the pump and that will tell you how big of a battery you need. Teslas have batteries ~ 50-70 kWh, so they could deliver that amount of power for perhaps 20-30 minutes.

IIRC there is a New Zealand group that flew a rocket named Electron which used battery operated pumps like this. They had to hot-swap and jettison the batteries during ascent. They had some switching problems if I recall but they were getting rockets toward space, if not into it. The technology is definitely possible.

I'll be over there calculating how much pressurizer I'd need, probably just to be shocked again.

I think the thing I love about being an engineer is calculating a big force/pressure/power and translating those numbers into a truly awesome mental picture of how much damage it could cause if something goes awry. I rather like horrifying my students by making them do this as well.

 

[jedidia]

Thanks for the pointers to the Electron, guys! I can't seem to find much details about their batteries, but it does sound impressive. I've thought of Tesla, obviously, but one of the things that makes me very hesitant is... Lithium on the Moon. Exists, but getting tons of it together is a steep ask, apart from the ton or so it would probably add to the dry mass.

In contrast, Getting 10 bars of pressure into those tanks seems to be possible with some 300 to 400 kg of additional oxygen. That would, in this cae, seem a much preferable option. However, there's an obvious problem with pressurizing a tank of liquid oxygen with gaseous oxygen. Can you even make that work? I imagine either the gaseous oxygen won't stay gaseous for very long (there goes the pressure), or the liquid oxygen won't stay liquid (which I'm not sure how bad it is...?)
Obviously, that additional oxygen would require additional tanks, but it should still beat out the mass of a pump and a battery by a long shot (and in the end, those tanks need to be filled with something, so even then I need some excess to evaporate).

I'm picturing something like a chamber/nozzle cooling cycle that heats up and evaporates liquid oxygen, then feeds it back to put more pressure in the tanks. Could you tell me, without going much into any grity details, whether such a thing is possible, or whether there are too many problems with it to shake a stick at?

As an alternative, the RCS will probably need to go the powdered Aluminium route, as having a solid fuel grain in each RCS thruster seems horribly troublesome. So I guess I could take a bit more of that along to also fuel a pump generator, but that would undoubtedly also mass in a lot more than the pressure solution in this case. If there's a chance that such a solution can be made to work at all.

 

[Thunder Chicken]

However, there's an obvious problem with pressurizing a tank of liquid oxygen with gaseous oxygen. Can you even make that work? I imagine either the gaseous oxygen won't stay gaseous for very long (there goes the pressure), or the liquid oxygen won't stay liquid (which I'm not sure how bad it is...?)

Having both phases in the tank at equilibrium is called a saturation state, and that could be useful as the saturation pressure is a direct function of temperature. You could attempt to run the tank at the saturation pressure of LOX, which varies with temperature according to the blue line in this diagram:

The trick is that you would need to keep the tank temperature below -150C or so to keep tank pressures manageable, or otherwise vent boil-off continuously which would be wasteful.

It's similar to how cans of so-called compressed air work. They are really cans of a saturated refrigerant (liquid and vapor phases coexist at equilibrium). As long as there is liquid and vapor in the tank, you can get a constant pressure and mass flow rate for as long as you push on the nozzle.

 

[Thunder Chicken]

I actually Googled the phase diagram for flourine to see if maybe that would provide a better saturation pressure / temperature combo. The first step to recovery is acknowledging that I have a problem...

 

[Urwumpe]

In contrast, Getting 10 bars of pressure into those tanks seems to be possible with some 300 to 400 kg of additional oxygen. That would, in this cae, seem a much preferable option. However, there's an obvious problem with pressurizing a tank of liquid oxygen with gaseous oxygen. Can you even make that work? I imagine either the gaseous oxygen won't stay gaseous for very long (there goes the pressure), or the liquid oxygen won't stay liquid (which I'm not sure how bad it is...?)
Obviously, that additional oxygen would require additional tanks, but it should still beat out the mass of a pump and a battery by a long shot (and in the end, those tanks need to be filled with something, so even then I need some excess to evaporate).

Depends. If your tank is small, the additional mass for thicker walls won't really kill you to keep 10 bars more in the tank. But for large tanks and in spaceflight, this gets a bit different.

Lets say, you want a burn time of merely 30 seconds, so your tank only needs to hold about 1500 kg of LOX or 1.65 m³ plus ullage volume, lets say 4% of keeping thermal expansion at full tanks under control, so around 1.7 m³ tank volume. Small enough for a spherical tank. Inner radius is just 74 cm, lets assume a rather normal aluminum alloy being used there.

Case	Wall thickness	Wall structural mass
250 kPa	1 mm	11.6 kg
1000 kPa	4 mm	67.9 kg

 

[jedidia]

so your tank only needs to hold about 1500 kg of LOX or 1.65 m³ plus ullage volume

Hmmm, ufortunately my tank volume for the LOX would be about 12.4m³ without ullage (though that would hardly all go into one tank), so I guess that would add quite a bit. Will have to run the numbers on how much.

The trick is that you would need to keep the tank temperature below -150C or so to keep tank pressures manageable, or otherwise vent boil-off continuously which would be wasteful.

Yeah, venting anything that doesn't absolutely need to is pretty much off the table in space flight. But good to know that there would be ways in which this could work (I don't really need to know how exactly, that would be going into way too much detail for my purposes. All I need is plausibility, not to build an actual spacecraft... ).

 

[jedidia]

Inner radius is just 74 cm, lets assume a rather normal aluminum alloy being used there.

Say, do you perchance have the equation handy with which you calculated this?

 

[Urwumpe]

Say, do you perchance have the equation handy with which you calculated this?

t = p * r / (s - 2*p)


p = pressure (Pa)
r = inner radius (m)
s = tensile strength (Pa)
t = wall thickness (m)

But thats only for spherical tanks, I would need to look for the other one if you want cylindric tanks.

 

[jedidia]

Spherical should be fine, I think. I like the looks of spherical tanks. Yes, I know that's not the most important thing...

After a quick back-of the envelope using 4 tanks and a tensile strength of 450 MPa (should be possible with an aluminium-titanium alloy, both available on the moon. Will probably have to throw in some imported ingredients, but hopefully not too much), the results look pretty encouraging that this won't blow up the mass to ridiculous proportions. Will have to dive into it a bit more when I'm off-work, but it looks more and more like pressurization by additional oxygen is indeed the most effective way to get my LOX to the combustion chamber and won't change the initial outline of the design by much. Great! Thanks everybody!

 

[Linguofreak]

Yes, but could you also get a material that withstands a combustion temperature of about 6 * 2800 K? Again, these are just rules of thumb derived from the idealized equations for characteristic velocity and assuming a constant ratio of specific heats. Of course, in reality, many more parameters would change at higher pressure.

Wait... The NASA experiment was only 2800 K chamber temperature? How did they get any specific impulse out of that? That's below the boiling point of Al2O3, so you'd get zero chamber pressure, right?

The stuff boils at 3250 K, so your chamber temp would have to reach at least that to have a viable engine to my understanding, but apparently it worked on the test stand???

 

[jedidia]

The NASA experiment was only 2800 K chamber temperature?

I'm not sure where that number is from, I couldn't find any mentions of chamber temperature in the paper.
P.s: It does seem to be the lower end of the combustion temperature of aluminium.

 

[Linguofreak]

I'm not sure where that number is from, I couldn't find any mentions of chamber temperature in the paper.
P.s: It does seem to be the lower end of the combustion temperature of aluminium.

I couldn't find it in the paper either. Urwumpe, could you clarify where you got this from?

In any case, at that temperature I'd expect an exhaust comprised mostly of glowing, slaggy smoke with almost no performance unless something else added to the fuel or oxidizer provides a filler gas to be the actual working fluid.

Maybe they were running super lean with extra O2 as the working fluid? That sounds likely to develop engine-rich exhaust problems, though...

 

[Urwumpe]

I couldn't find it in the paper either. Urwumpe, could you clarify where you got this from?

Its actually in the paper, right in the beginning. Also note the abstract of the paper, while the actual download might be again impossible due to server issues:

Aluminum combined with oxygen has been proposed as a potential lunar in situ propellant for ascent/descent and return missions for future lunar exploration. Engine concepts proposed to use this propellant have not previously been demonstrated, and the impact on performance from combustion and two-phase flow losses could only be estimated. Therefore, combustion tests were performed for aluminum and aluminum/magnesium alloy powders with oxygen in subscale heat-sink rocket engine hardware. The metal powder was pneumatically injected, with a small amount of nitrogen, through the center orifice of a single element O-F-O triplet injector. Gaseous oxygen impinged on the fuel stream. Hot-fire tests of aluminum/oxygen were performed over a mixture ratio range of 0.5 to 3.0, and at a chamber pressure of approximately 480 kPa (70 psia). The theoretical performance of the propellants was analyzed over a mixture ratio range of 0.5 to 5.0. In the theoretical predictions the ideal one-dimensional equilibrium rocket performance was reduced by loss mechanisms including finite rate kinetics, two-dimensional divergence losses, and boundary layer losses. Lower than predicted characteristic velocity and specific impulse performance efficiencies were achieved in the hot-fire tests, and this was attributed to poor mixing of the propellants and two-phase flow effects. Several tests with aluminum/9.8 percent magnesium alloy powder did not indicate any advantage over the pure aluminum fuel

Pay attention to the mixture ratio and also the "two-phase flow effects". It did not boil at all.

 

[Linguofreak]

Its actually in the paper, right in the beginning.

I was specifically asking about the chamber temperature of 2800 K you quoted, which I don't see in the paper.

Pay attention to the mixture ratio and also the "two-phase flow effects". It did not boil at all.

I don't know what the stoichiometry for the reaction is and haven't has time to look it up/figure it out. How far rich/lean of stoichiometry are they going?

I guess they'll have at least some gas in the exhaust flow at any mixture ratio given that they note incomplete combustion, but having your working fluid be O2 at 2800K doesn't sound like it's good for engine life.

I wonder how water/aluminum would do? Your oxidizer doesn't have to be cryogenic, you have pure hydrogen as your working fluid assuming stoichiometric burning and complete combustion, and water won't be as nasty too the engine if you run lean (and has a lower molecular mass than O2).

EDIT: Apparently, tests have been done on an aluminum/ice solid propellant ("ALICE").

 

[Urwumpe]

OK, is not mentioned in this paper at all, as I thought to remember, but another one puts 2800K at the lower end of aluminum powder and oxygen combustion.

https://pdf.sciencedirectassets.com/271494/1-s2.0-S0950423020X00064/1-s2.0-S095042302030557X/am.pdf?X-Amz-Security-Token=IQoJb3JpZ2luX2VjEJ7%2F%2F%2F%2F%2F%2F%2F%2F%2F%2FwEaCXVzLWVhc3QtMSJIMEYCIQDSZrGNhjltoUeZpuDCDWP2UYcVUjktLCqPalv3EGXz0wIhAM%2Fn0xtDeoCvoYq1huZZQ92zqXQ5PWKSgpPUYmNSzUm2KrIFCBYQBRoMMDU5MDAzNTQ2ODY1IgxuuIwDytFiT5v2gekqjwUAbBpGbmr9BpA7MgkIZKqBgXvHavamhh3%2BLiky%2Bf6njbrryKPToLqk1C%2FVSSwU6gz2%2BasPEiqHPQ8m7Ug8MezcazV3bjcGbAhP8w4Es1X51pRxGFAWxe8Rla%2B2s4KfnGnGPYCZ%2F7eegwX9pmphIiLr0u0XjTvr%2FDHCoRiOEwfaO1NaJYFLJPG7iXPFmnk8JEoXKpcnUZ7a8zZxv0B6EQonBAZQW8iW2ymQxBoDmP2JmHnyuCDRNB%2BxJ3778AFnE0mUyOcQIxhtTbP%2BOvCv0x9O5R7zhDXNOuKQOA3fpwSoqXDYhjj5WxN%2BLozkeQZdh2BlLdLN5%2BXh8iw5c92MWZeJt6WxtdlbcyTaL5ZUqTHzNiyEWnUmNWN7RL%2F%2BCZoSBdPZlVRwiJl5S9vN%2B5HT0Ny1YZxACPyNXg9XkDn3qBjOQP3IIxkwlC5LzEsSltN3MZGWcrXHxecqKOjUJG3cCggGNfNUPB505YfdvYl2Try9a3a5GFPAJ9k%2BNvggfu0VmVYl2js518NYkCUuENJ8%2BX1oZisssq1ADpVIr6AsHF8nqlEHMrFaC1gsFa26E9NF3avWBnhWyvYuBIMdgvjQPHKkEEnK%2FxbOk%2FZdyHZM2fp3RzRXlVy%2FHVxTeeNi9HaESlE4Uv8b7Wmw4XVVPcycwBa3U95PxDgZxHrKaA5NfliaSz0BJ5FyLjEUuP5LnVHR92Sp%2BJ5syf%2F7kWsL2zHH4eYJP2eZadyCS9pSyNHXEvX%2FUs6MrNOYJW2nu%2Bb%2BYGtrFkDls3Y5KoHMayhjns%2BeNBj5r0h79lqIuh%2FzwQ%2BKWbgFF%2F20mDJG86EhLTOY6IVd0%2FqbC1Mdr2s%2Bnfo9pbXDFrrP6Wua5Mcs9iKbaIbqnf0wML%2Bzv78GOrABkO9ZnKPKHigfUTybEZexiPCO7cMaE9O4YZUCaR%2FFzHW14AwAfc5OM2H72tk99R4YzkE9KSPEutsXcEL5Lc%2BYkZ2J3IEKYUb3AIKZ%2FSs4mnfNPDJWWL%2Fvx03fLq5zMciA0IxXZ48SHk%2Fh04TGxVU8X8fYE6o4nuEzjpcDgcXbBX4Bwp21290hndG03Ep4ZPbnI0FKfjv0kStOJEWSd08ZTye5McGB%2BfFtO2hFw5HNMrI%3D&X-Amz-Algorithm=AWS4-HMAC-SHA256&X-Amz-Date=20250404T134431Z&X-Amz-SignedHeaders=host&X-Amz-Expires=300&X-Amz-Credential=ASIAQ3PHCVTYSETGVQZE%2F20250404%2Fus-east-1%2Fs3%2Faws4_request&X-Amz-Signature=e91090a3d5d2b8793712f0ded86d7cf3d001678ac7de42a1820fa84d21395855&hash=7b710e2d908ae4eb200cf52d2ba981b0d7f8f623231aa7543ddceaabaad95bcf&host=68042c943591013ac2b2430a89b270f6af2c76d8dfd086a07176afe7c76c2c61&pii=S095042302030557X&tid=pdf-eac80127-ad9d-4e80-bafc-02af807d4598&sid=12c91e48538d3841733b5810db55374bb322gxrqb&type=client

Sorry!!!!

 

[Buck Rogers]

This might be useful?
Estimation of Launch Vehicle Propellant Tank Structural Weight using Simplified Beam Approximation. -Georgia Inst. of Technology

Link to all the files:

Dave Akin's Web Site

 

[Arvil]

I wanna see someone type that into a search bar manually without copy/paste