Rocketlab's Electron uses an electric pump combustion cycle with no preburner so all fuel is fed into the combustion chamber rather than wasted in a preburner/staged combustion-powered turbopump that requires complex engineering and finicky systems so my question is why do/are larger new rockets not using electric pump cycles? With improving battery and electric motor technology I would think that the fuel efficiency, throttle ability, and ease of engineering would be desirable especially with reusability being a concern now. In something like starship that will be coasting through interplanetary space, you could use the methane and oxygen boil off to power a fuel cell that would in return recharge turbopump batteries or you know you could have solar panels powered by the sun. Is weight really the only factor, I mean a super complex turbopump with all the crazy pressure budget compatibility like in the Raptor can't be that much lighter?

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    $\begingroup$ Re the SSME "its high-pressure fuel turbopump alone delivered as much horsepower as 28 locomotives, " That's a lot of batteries. ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/20120001539.pdf So yes, "a super complex turbopump with all the crazy pressure budget compatibility " can be a lot lighter. $\endgroup$ – Organic Marble May 6 at 18:38
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    $\begingroup$ Wait, so you are telling me there really is a difference between a 5,000lb engine and an SSME?? Hahaha, thanks for the info! $\endgroup$ – YuccaWorks May 6 at 18:42
  • $\begingroup$ A good throttle ability requires not only to throttle the pumps but also the nozzle of the combustion chamber. At 50 % the neck width should be much smaller. When using the same neck width as for 100% chamber pressure at 50 % would be much too low. A low pressure would decrease exhaust velocity and efficieny. $\endgroup$ – Uwe May 6 at 18:56
  • $\begingroup$ Makes sense. Thanks! $\endgroup$ – YuccaWorks May 6 at 19:00
  • $\begingroup$ Batteries are heavy. And pressure is irrelevant because even if you electrify it you only replace the turbo with a electric motor but the pump side is unchanged. $\endgroup$ – user3528438 May 6 at 20:47

Ultimately the pump, whether electric or combustion-turbine driven, needs a certain amount of total energy input to do its job.

Combustion reactions both deliver more total energy per mass than batteries, and deliver it faster. From an article on gas versus electric cars:

Stored energy in fuel is considerable: gasoline is the champion at 47.5 MJ/kg and 34.6 MJ/liter; ... A lithium-ion battery pack has about 0.3 MJ/kg and about 0.4 MJ/liter (Chevy VOLT). Gasoline thus has about 100 times the energy density of a lithium-ion battery. This difference in energy density is partially mitigated by the very high efficiency of an electric motor in converting energy stored in the battery to making the car move: it is typically 60-80 percent efficient. The efficiency of an internal combustion engine in converting the energy stored in gasoline to making the car move is typically 15 percent (EPA 2012).

So the mass ratios are significantly in favor of combustion for cars -- at least 35:1 including the efficiency factor (but this article is from 2012; I believe battery packs have gotten much better since then). As user3528438 points out, this doesn't account for the turbopump's oxidizer consumption (cars get their oxidizer from the air instead of carrying it), so it's probably somewhere in the 10-15:1 range instead.

If the electric pump can be made much smaller and lighter than the equivalent-power turbopump, it might close the gap a little more, but I doubt that it's a mass win for the Electron.

Your suggestion of recharging the batteries during an interplanetary coast, so they need to only be sized for the more demanding injection phase, rather than injection + descent, is interesting, though, and might improve the calculus a little bit.

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    $\begingroup$ +1 There was a project to replace the hydrazine powered APUs on the shuttle with electric pumps and batteries. They just couldn't make it work mass-wise. $\endgroup$ – Organic Marble May 6 at 19:02
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    $\begingroup$ This is not completely fair because first of all you didn't include the weight of the oxidizer in the energy density. Also not all rocket engines are closed cycle so those with open cycle will have much lower energy density when they dump the fuel/oxidizer exhaust. $\endgroup$ – user3528438 May 6 at 20:49
  • $\begingroup$ Your first point is legit. Turbopump combustion is deliberately quite fuel-rich to keep temperature down, so let's assume the mass ratio is around 1:1 (2.4:1 lox:kerosene is typical for the main combustion chamber) -- that cuts the turbopump's advantage in half. However, my comparison is equal-basis assuming open cycle and zero contribution to thrust, which is not typical. If the turbopump exhaust is pointed downwards or ducted into the nozzle, there's some small increase in thrust. Closed cycle will come out further ahead. $\endgroup$ – Russell Borogove May 6 at 21:29
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    $\begingroup$ It needs to be extremely rich or lean or you over heat the turbine. If your preburner has the same near-ideal mixture ratio as main chamber it will just be as hot, do you think that's a good idea? $\endgroup$ – user3528438 May 6 at 23:48
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    $\begingroup$ @RussellBorogove if you're talking about the gas generator on the F-1, its mixture ratio was 0.416. pdf-archive.com/2016/10/21/rocketdyne-f1-engine-manual/… page 1-6 $\endgroup$ – Organic Marble May 7 at 2:17

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