Going off memory, and I’m too lazy to verify any of this, but here goes (the numbers are rough, but decent):
The first stage of the Saturn V boosted it to roughly 40 miles high, and 6000 mph.
The second stage boosted it to a little over 100 miles, and a little over 15000 mph.
The third stage then lofted it into orbit and was shut off, while they orbited to check out everything before hitting the gas again for trans-lunar insertion.
Was it just impractical for weight considerations to have two stages to orbit, then the third stage to fire only once for TLI?
I think it was done partly for flexibility and partly to build on work that was already done. The upper stages were very similar to those done on earlier Saturn rockets.
A two-stage Saturn V could get you into an earth orbit. Skylab was launched from one of these. If you needed to go further (like say to the moon) then a three stage version could be used.
The S-IVB (3rd stage of the Saturn V) as also used in the Saturn IB stack as the second stage with some minor modifications. Seemingly, the modifications were so minor they felt no need to call one version the S-IVC. So they found a way to use one stage in two different stacks and types of missions. And perhaps the first two stages of the Saturn V might have had enough power to boost Skylab into orbit but not enough to lift the CSM, LM, and S-IVB into orbit?
IIRC, Apollo XVII was the heaviest load lifted until sometime in the 1990s, by the big Russian booster. So the moon missions were considerably heavier payloads. LEO is LEO and it doesn’t take much delta-vee to get out of Earth orbit regardless of stable orbit height. So, the answer is “greater lift capacity.”
Saturn V’s. How can you not like something that has 75,000 HP fuel pumps? “Ah’ve burned more alcohol in sixty seconds than you’ll ever serve across this lousy bar.”
I remember thumbing through a book of thermodynamics data back in college. One page folded out into a long horizontal graph of the performance of the Saturn V, with altitude, speed, and weight (and probably a few other parameters) as a function of time. The weight data was staggering; that thing burned off a few hundred thousand pounds of fuel in the first couple minutes.
Tossed this out to my son - an aero engineer w/ ULA. He says:
Several reasons, mostly Dry Mass Fraction - the % of the rocket that is not fuel, i.e. tankage, and payload. You need a high dmf to get to orbit. So high, in fact, that no rocket with current technology can do it with a single stage. When rockets dump tankage halfway up, the fraction of payload (dry mass) continuing to orbit goes way down.
Today there are a couple of proposals to get to the moon and back w/ 2 stages. Saturn V needed 3 stages, lunar orbit rendezvous, and a staged lander because their rockets were not as light as today, engines were less capable, and they had a huge payload.
Sorry if I garbled anything in relating his texts, but that is the gist of how 1 rocket scientist responded to the OP. Let me know if you have any more questions and I’ll pass them on.
Also worth mentioning - the Saturn V was not custom built for a lunar landing. It was already broadly specified in design before the Apollo programme. It was intended to be a general purpose heavy lifter that could be used to power the imagined US space efforts long after Apollo. As it was the funding was never there, and Skylab was the last flight.
Before lunar rendezvous was established as the preferred mission profile for Apollo there were profiles that needed two Saturn Vs to perform a mission.
I remember from the time of the Space Race that the Saturn V design parameters had been decided well before we decided to go to the moon. Also something about the fuel tanks in the first two stages being tested and known as reliable designs. A two stage rocket would have required completely new tank designs.
Rated payload of a Saturn V counting on using the S-IVB stage: 118 metric tons (nothing else ever since has even come remotely close, not even Energya)
Mass of a fully fueled S-IVB: 119 metric tons (90% of it fuel/oxidizer); mass of the full 3 modules of Apollo spacecraft in the later missions: 45 metric tons.
Duration of burn to Parking Orbit 2.5 minutes, duration of TLI burn 6 minutes. Call it one fourth of the fuel used for orbital insertion and three fourths for TLI.
Mass in orbit of the S-IVB+Apollo stack for Apollo 15 (one of the heavier missions, with the moon buggy and the works): 141 metric tons, after the orbital insertion burn. So something like 25+ tons of fuel used for LEO insertion.
This isn’t strictly true. Although the ‘Saturn family’ of vehicles (with designations from C-1 to C-5) was generally conceived by Von Braun’s ABMA program, the Saturn V was purpose designed to support the J-class missions, and the system requirements reflect that. While Von Braun had greater plans requiring much larger rockets (dubbed the Nova family) the Saturn family was explicitly limited the capability required for the lunar landing and test flights to support it. As far as NASA of the pre-Apollo era was concerned, there was no “beyond Apollo”, and despite the best efforts of both contractors and groups within NASA promoting a variety of mission concepts beyond the J-class missions using modifications of both Apollo and Gemini systems, there were never any firm plans beyond the very hobbled Apollo Application Program (resulting in the Skylab flights and the Apollo/Soyuz Test Program) that resulted in any contractual developments.
It is true that the Saturn V would not support the direct ascent profile for an Apollo-class mission, and the Earth Orbit Rendezvous missions would have required multiple launches (depending on configuration). But the study on mission configuration fed directly into the system requirements analysis that defined the Saturn V specifications.
The S-I first stage of the Saturn I and Saturn Ib used Redstone and Jupiter tanks grouped together. However, the S-IC (first stage of the Saturn V) was purpose built by The Boeing Company to use five F-1 engines (versus the eight H-1 engines used by the S-I stage).
The reason that three stages were used in the Saturn V is essentially laid out by Dinsdale; the mass fraction required by the design of the stages and performance of the engines was such that the third stage–also used as the trans-lunar injection stage–was necessary to obtain the required orbit. With better mass fractions allowable by the use of lighter and stronger materials (the aluminum-lithium alloys used in the first stage tank) and greater performance of modern engines, the use of an intermediate stage between the ‘ground level’ and exoatmospheric performance. (The advantage of the compactness of hydrocarbon fuels and reduction of the dead weight of tankages outweighs the propulsive efficiency of less dense liquid hydrogen fuels in a vacuum environment.)
A reusable single stage to orbit (RSSTO) vehicle is the ultimate holy grail of ground-to-space transport. Such a vehicle is, at least theoretically, within reach, and one rejected proposal for the Space Transportation System (the Chrysler Aerospace SERV which gave us the decidedly non-efficient Space Shuttle) would have provided significant payload capacity to low Earth orbit (LEO), albeit insufficient to support Moon or interplanetary missions. However, for the reasons that Dinsdale stated, multi-stage vehicles offer greater propulsive efficiency, and given the primitive state of the art in the early 'Sixties the three stage Saturn V represented the best configuration for a moon shot. However, it should be noted that the Titan II GLV, used for all of the Gemini missions to LEO, used two stages both powered by nitrogen tetroxide and Aerozine 50, were sufficient for those missions.
Here’s a simplified version of what I thinkDinsdale is saying:
As the fuel tank empties, the now-empty portion of that tank is just dead weight, making the rest of the ride less efficient in terms of fuel use. From one perspective, it might be a great idea to have a dozen (or a hundred!) fuel tanks, and jettison each one as it empties, so that you don’t have to carry them further up. Stay lean and mean!
But really, that’s overdoing it, because the systems to hold all those tanks would add too much weight on their own. So the designers try to find a happy medium somewhere. Gotta have at least two stages, and for the Saturn V they settled on three. (Personally, I’d argue that when the Apollo is put on top of the Saturn V, the Service Module counts as the fourth stage.)
Minor nitpick by a daddy very proud of his son, but any remarks in this thread attributed to Dinsdale came straight from Dinsdale’s son.
The kid pretty much lives and breathes manned space flight. Some other info from from his texts last night which might add to the conversation:
Just thought I’d toss this out there. My kid isn’t THE expert in such things, but he sure devotes a lot of time and thought to this stuff, so it might add to the discussion.
[QUOTE=Dinsdale’s son]
Currently almost all launch vehicles ar 2 stage to orbit (TSTO). . . .
[/QUOTE]
Which was actually the question I was trying to pose. Not “Why did it take three stages to get to the Moon,” but “Why did it take three stages to loft into orbit?” (Well, technically, about 2¼.) From the responses it seems that the existing technology was pretty reliable, and it would have been impractical to redo the second stage to give it enough oomph to attain orbit.
The Service Module is part of the payload (along with the Command Module and for some missions the Lunar Module and Lunar Roving Vehicle) and therefore not considered part of the booster configuration.
It should be pointed out that it isn’t just the weight of the tankage but also the engines and associated plumbing, interstates, and other hardware that makes up the dead (non-propellant) weight, and the propulsive requirements change with altitude and speed. At ground level, very high thrust is required to get the vehicle to go anywhere, to the point that less efficient propellants are used, e.g. why the Saturn first stage (S-I) and most other liquid propellant boosters today used RP-1 (high grade kerosene) and liquid oxygen, or why solid propellant boosters were used on the STS, Delta IV medium, Titan IV, and many others. Despite the inherent low specific impulse (I[SUB]sp[/SUB]) of the propellants due to the high molecular mass of the combustion products, the high thrust of the boosters and density of the fuel (therefore requiring less tankage) is of greater advantage than pure propulsive efficiency, especially in dense atmosphere where the ambient back pressure will prevent the plume from expanding efficiently. Rocket engine nozzles designed for sea level operation are also shorter and generally designed to operate for only a few minutes.
Upper stages which operate in low pressure or vacuum conditions and are designed to sustain and slowly boost the payload into orbit (sometimes with multiple burns) need less thrust, will have larger (and sometimes extendable) nozzles to capture the full expansion of the plume, and will be designed to operate for longer durations.
An example of a hybrid of this is the original SM-65 ‘Atlas’ (originally designed and briefly deployed as the first ICBM in the US arsenal). The Atlas was technically a single stage vehicle with all three engines ignited on the ground; however, once it gained sufficient altitude and speed, the two outside engines were dropped and the “sustainer” engine (which was of a slightly different configuration) propelled it through the remainder of powered flight. This was not done because it was the most propulsively efficient arrangement, but because of concerns about staging and starting an engine in mid-flight which, in retrospect, were very valid, as a significant number of failures have and continue to occur during the staging operation.
No, this is not correct. There was no existing second stage to have to redo; the Saturn V was purpose built for the Moon landing program. Note that other versions of the Saturn family which actually flew (Saturn I, Saturn Ib, and Saturn INT-21) were two stage vehicles, as was the Titan II GLV which flew the Gemini missions. The reason is that the specific mission profile of the F-class missions and beyond required a restartable third stage to inject into lunar orbit, and it made more in a heavy boost vehicle to allocate the propulsive impulse to achieve a pre-trans-lunar injection to the third stage as it was already being designed to operate in pure vacuum conditions as opposed to the S-II, which was optimized to operate at high altitude endoatmospheric conditions. Follow-on configurations of the Saturn family proposed for both Lunar and interplanetary missions would have included super heavy lift vehicles which used two main stages (an uprated S-I and various configurations of the S-IVB) along with solid boosters as Stage 0.
Dinsdale, your son is correct that many newly designed systems for orbital launch use only two stages for primary boost; however, vehicles with three or more stages certainly exist (Minotaur, Indian GSLV, most of the Chinese Long March family). This is a tradeoff between the greater simplicity and reliability of two stages versus efficiency of three stages. It should be noted that the Sea Dragon concept, which was a simplified ultra heavy lift vehicle (also known as a Big Dumb Booster) proposed by Bob Truax (which traded efficiency for shear payload mass) was only two stages and was judged to be more cost effective and potentially as reliable as existing launchers.
An expendable single stage to orbit (SSTO) vehicle is, despite claims to the contrary from many quarters, entirely practicable even without using exotic materials or altitude compensating nozzles. However, the payload to gross liftoff weight ratio is not adequate to justify the costs for a heavy lift vehicle, and the costs for a personnel transporter are significantly better than a two stage booster. A reusable SSTO would be desirable for flying personnel or small payloads into orbit (hence, why it was pursued by the US Air Force, the Strategic Defense Initiative Office, JAXA, and others) but the required technology to make the entire vehicle capable of withstanding reentry and being reused (as opposed to just a capsule or small lifting body shuttle) is just beyond the current state of the art. The previously mentioned Chrysler SERV would have provided just that, but while a review of the studies indicated it was feasible, the sensitivity to weight and the performance of the then-untested aerospike nozzle, along with the greater cost versus the parallel stage solid booster and liquid engine ‘Shuttle’ meant it was a greater risk than NASA was willing to take.
One other thing: I inserted a word into your quote because I was wondering if that was an omission on your part, or did you mean it exactly as it was written.
I’ve always like the Big Dumb Booster concept. Who cares if the bottom stage is way bigger than a high performance model? Its the price that counts. Stuff that’s more a shipyard quality build is cheaper than cutting edge aerospace stuff (save that for the upper stages where it makes sense.
What’s you opinion on BDM Stranger?
I think its a shame its never been tried (well they did a small scale demo one and tested it on the ground IIRC).
And I take it the real world numbers for just shedding tankage just don’t work out?
It seems NASA had another shot at this concept with the X-33/VentureStar? Would they have pulled it off if they didn’t get so ambitious with new technologies like composite cryogenic tanks?
When the goal is simply Earth orbit (Apollo 9), I agree that the SM and LM were payload. But I figure that if you want to go to the moon (where you need the SM for course corrections), and certainly if you want to get BACK from the moon, surely it’s not mere payload!
Maybe I don’t understand the word “stage” correctly.
