The images of the engine clustering in the Rocket Labs's Electron rocket gives an indication that there are independent drive pumps for individual engines; I am not sure of it. I am not aware of the plumbing of the Falcon 9 too!

Does having independent pumps for every engine in a clustered configuration have any advantage over a common pump configuration that splits the propellant into the $n$-engines?

Does every clustering configuration has the same pump configuration or are there any tradeoffs that are considered to adopt common/independent pump arrangement?

  • $\begingroup$ Pure speculation: 1 pump fails? 1 engine fails. Versus: 1 pump fails? Rocket stops accelerating. In addition doesn't having a pump on each individual engine allow for individual up/downthrottling on each engine? space.stackexchange.com/questions/7841/… $\endgroup$ Jul 19 '18 at 12:17
  • $\begingroup$ Yeah, redundancy is one of the main concern! Btw, what is the practice in clustering? say for in SSME, Saturn V, Falcon, Soyuz, etc $\endgroup$ Jul 19 '18 at 12:19
  • $\begingroup$ I'd read the link I added in an edit, seems they have good info on the Falcon 9 attached there in terms of why they decided on 9 engines. $\endgroup$ Jul 19 '18 at 12:20
  • $\begingroup$ @MagicOctopusUrn - but wouldnt a single pump make things easier? Also is the same independent pump configuration used in all engine clusters? $\endgroup$ Jul 19 '18 at 12:25
  • $\begingroup$ Easier is usually not the goal. Redundancy, reliability and safety are what they're going for. I'm guessing that, while it would make the engineering easier, it would make the overall risk higher. I don't know answers to the technical questions, but there are great sources on Falcon 9 everywhere. $\endgroup$ Jul 19 '18 at 12:27

In this post, I'm talking about engines slightly abstractly, using "pump" to mean the entire pumping complex or "powerhead", which might be made of multiple actual turbopumps, and "combustion chamber" to mean the main combustion chamber, not any preburners or gas generators.

Most often you'll see one pump per combustion chamber and nozzle.

Some large engines, mostly Russian, split the combustion chamber: large chambers have more issues with combustion stability, while pumps appear to scale up more straightforwardly. The Soyuz's RD-107/108 engine comprises one turbopump and four combustion chambers + nozzles, for instance. The single-pump, four-chamber RD-170 used on Energia's liquid fueled boosters has 2-chamber (RD-180, used on Atlas V) and single-chamber (RD-191, used on Angara) derivatives using smaller pumps; these must have been very straightforward to design since the chambers and nozzles were already proven.

The first-generation Atlas rocket had a two-chamber booster stage; the design actually switched from a common turbopump to separate turbopumps and back to a common pump in various versions, which suggests that there isn't a huge advantage to one configuration over the other. The gas generator component was shared on all versions, which further complicates the counting.

The F-1 engine used on the Saturn V was large enough to run into these combustion stability problems, but they kept redesigning the injector and upper portion of combustion chamber until it worked, rather than redesign it as a multiple-chamber engine.

Saturn IB's H-1, Shuttle's SSME, Falcon's Merlin and the Electron's Rutherford engines are all one-pump-one-chamber engines used in clusters.

The pump's mass flow rate is essentially proportional to thrust. For a given pump rate, the engine is generally lighter if you're using a single chamber instead of multiple chambers. So from a thrust-to-weight standpoint, the ideal would be one big engine per stage: a single pump delivering the needed total thrust and a single combustion chamber.

In practice, there are several reasons to consider clustering smaller engines to reach the same total thrust: engine-out capability; throttling by turning off individual engines, and avoiding the engineering and logistics problems found in building gigantic single engines.

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    $\begingroup$ Flip your mental model from "the combustion chamber is the engine and pumps might be shared" to "the turbopump is the engine and the chambers might be split", then re-read my last two paragraphs. $\endgroup$ Jul 19 '18 at 16:19
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    $\begingroup$ Each engine had four separate turbomachines. The high pressure LOX turbomachine had 2 pumps on it. That's 5 pumps. There was one main chamber and 2 preburners per engine. $\endgroup$ Jul 19 '18 at 16:23
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    $\begingroup$ Thrust-to-weight, reliability, throttling options, and ease of development. $\endgroup$ Jul 19 '18 at 17:05
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    $\begingroup$ I doubt there's any book that discusses this specific case in detail. It would be a mechanical/thermal engineering exercise to design two equivalent thrust engines, one with single chamber configuration and one with multiple. Sutton plus an appropriate engineering degree plus a detailed mass breakdown of an existing engine would probably be enough -- I've only got one of those three at hand. $\endgroup$ Jul 19 '18 at 17:52
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    $\begingroup$ um, I moved on 22 hours ago. $\endgroup$ Jul 20 '18 at 14:28

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