I've heard that the size of Saturn V was pretty much as big as could've been built in the 1960s. And the Soviet N1 did basically fail because it was too large for their infrastructure and engine size at the time. Generally, there's an economic advantage with larger scale, given enough demand for the-bigger-the-better one piece payloads, which I want us to fictionally assume here. What are the technical challenges that limit how large a chemical launcher, from the surface of Earth to orbit, could be built?
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1$\begingroup$ The Saturn V was as big as necessary for the moon landing. In the early 1960s some larger versions were planned under the name Nova. Saturn V used 5 F-1 engines for the first stage, there were Nova versions with 8 F-1. Nova development was cancelled 1964. $\endgroup$– UweMar 23, 2017 at 17:46
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5$\begingroup$ FWIW, the N1 failed in large part because the Soviets couldn't manage the coordination between lots of engines, and harmonics kept blowing things up. Modern computers and electronic controls would have little trouble. $\endgroup$– chrylis -cautiouslyoptimistic-Mar 23, 2017 at 20:20
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1$\begingroup$ "I've heard that the size of Saturn V was pretty much as big as could've been built in the 1960s" -- do you have a citation for that? It's true in the relatively narrow sense that the development program for a larger rocket taken on in 1961 probably wouldn't have been completed by 1970, but not due to purely technological limitations. $\endgroup$– Russell BorogoveMar 23, 2017 at 20:47
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$\begingroup$ Why would you launch large one-piece payloads, rather than multiple loads that would be assembled in orbit? I'd think it'd be a lot cheaper to use 20 commodity, reusable SpaceX launchers than to design & build 1 rocket to lift 20X as much payload. Plus one failure doesn't doom the entire project - you just launch another sub-assembly. $\endgroup$– jamesqfMar 23, 2017 at 22:06
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1$\begingroup$ Many management costs scale less than linearly with the size of the launch, and on-orbit assembly comes with costs as well. One Sea Dragon at $2.3 billion per launch in today's money would be competitive with 20-30 Falcon 9s. $\endgroup$– Russell BorogoveMar 23, 2017 at 23:14
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4 Answers
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SpaceX is proposing launch vehicle, as a first stage, and a second stage that would transit to Mars. Interplanetary Space Transport (ITS).
The final size is not really confirmed, and they should actually build it before we compare it, but using chemical rockets, it is quite a bit larger than a Saturn V.
The Saturn V used 5 F-1 engines with around 1.5 million lbs of thrust. The ITS first stage is planned to use 42 (!!!!) Raptor engines for a total of 29 million lbs of thrust.
Clearly larger than a Saturn V is in the realm of possible.
Large engines are hard, as the F-1 engine development showed, and as the Soviet experience that lead to 30 NK-15/NK-33 engines for the N-1 showed.
SpaceX has taken a middle of the road approach, but closer to the N-1 approach with fairly large numbers of medium sized engines.
Moving the booster around is quite tricky. NASA solved it by barging the large components. US Rail does not support such large objects. Soviet rail seems like it was able to handle larger as most of their components were moved by rail.
SpaceX plans to solve it by flying the vehicle around. (Or at least manufacturing near the launch site, and after that, land and relaunch mostly in place).
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1$\begingroup$ ITS is only slightly taller and fatter than Saturn V, still it can (hopefully) launch so many more times payload to orbit? I'm sure there's plenty of speculations about that around, but in principle, don't launchers with chemical engines need to be larger in proportion to their capabilities? $\endgroup$ Mar 23, 2017 at 14:47
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$\begingroup$ Methane/LOX vs RP-1/LOX so fairly higher Isp even on first stage. Quite a bit bulkier. And even then, for Mars missions need 5 refuelling launches for the upper stage. $\endgroup$– geoffcMar 23, 2017 at 14:50
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4$\begingroup$ The relative visual sizes are deceptive; the Saturn V 2nd and 3rd stages are full of hydrolox, with a bulk density of about 0.36, while the ITS methalox propellant has a density of about 0.8 -- it's thus a much more massive launcher. $\endgroup$ Mar 23, 2017 at 18:18
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$\begingroup$ Also, the side-by side images may be misleading. Remember that volume goes up as the square of the diameter. $\endgroup$– jamesqfMar 25, 2017 at 21:20
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It's generally thought that the Sea Dragon design was technically feasible; this would be 18000 tons (6 times the mass of Saturn V) at liftoff and deliver 500 tons to LEO, comparable to the SpaceX ITS in expendable mode. Mass fraction and specific impulse would be worse than Saturn V but economy of scale would make it more cost effective per payload ton. It was never produced because the demand for such large payloads never arose, of course.
I don't know of any specific hard technical limitation to going bigger than Sea Dragon, though mass fraction probably gets poorer and logistic issues get hairier as you embiggen.
answered Mar 23, 2017 at 17:17
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$\begingroup$ For the record: ITS was said to have a capacity of 550 tons (possibly short tons, which would be about 500 metric) to LEO if used in expendable configuration. Since Sea Dragon wasn't supposed to be reusable, that's a fairer comparison... $\endgroup$ Mar 23, 2017 at 23:45
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$\begingroup$ Although Sea Dragon was supposed to be partially reusable with parachute recovery of first stage. $\endgroup$ Mar 24, 2017 at 2:12
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$\begingroup$ @RussellBorogove that is somewhat unlikely to have worked out. Parachutes don't scale well to the sort of size that would be necessary for the Sea Dragon first stage. $\endgroup$– LelielMar 24, 2017 at 2:51
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$\begingroup$ Sorry, I misspoke, it's not exactly parachute. The stage would be largely a big empty tank, with consequently low terminal velocity, and an inflatable drag flare further reduces impact velocity to a "mere" 90 m/sec. 55 tons of structure is allocated to the flare and recovery equipment, and another 20 tons of structural reinforcement relative to the expendable version. $\endgroup$ Mar 24, 2017 at 3:28
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Engines haven't changed much, so most improvements to size would likely come from the use of lightweight composite materials. The performance of these isn't as much better than aluminum alloys as is sometimes imagined though, so I doubt that would improve the situation a whole lot.
Nuclear thermal launch would improve payload fractions. The ability to use purely hydrogen propellant alone would dramatically improve exhaust velocity and thus the payload fraction of the rocket. This is essentially the only benefit, as in all such cases the temperatures and pressures are limited by the properties of chamber materials. It's quite a benefit though. The logarithmic term in the rocket equation makes high exhaust velocity extremely desirable. The key capability tradeoffs of course come from the high cost of nuclear systems and safety issues. It is worth noting though that no practical research has been in this area for several decades, so it's not clear what thrust to weight ratios could be achieved today.
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$\begingroup$ But how is the payload fraction a limit for launcher size? I edited my question to remove the distraction of nuclear propulsion, that was a side track, I'm sorry. $\endgroup$ Mar 23, 2017 at 13:12
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The major technical challenges of up-sizing rocket boosters, as I see it, are in two major categories: weight, and risk mitigation.
Weight
The most profitable business is weight reduction. There is money to be made on developing future materials (metals, and fuels) and structural design. These areas are relevant because the rate at which the size of the spacecraft increases, currently, will match and overtake the power efficiency very fast. Which is why we see the SpaceX Interplanetary Transportation System as three smaller craft taped together for added boosters.
Risk mitigation
Therefore the best thing for weight, multiple layered systems, is also the best thing for mitigation of risk for these "glorified explosions". Realize that not only are you controlling an explosion to propel these things into space, you are also preventing a fuel supply from igniting mid-trajectory. This is a very expensive cargo to be taking from point A to point B, and having a lesser amount of fuel mid-flight and capital loss to control is always prevalent.
