Big Dumb Boosters- and why we're overthinking this whole rocketry business

[magnemoe]

This is nonsense. It barely hits 200m.

During test flights where its not reason to go high and far as you focus is accurate landing.

Anyway Morpheus is an prototype of an moon lander, the software and control can be used on other bodies too, main point is that the lander can find its own safe landing spot.

Methane and oxygen should keep for an trip to moon but they need storable propellant for Mars or other places.

Turbo pumps are most useful for larger stages not small stuff where the extra gas don't weight much, I read another article there they wanted to use oxygen, hydrogen or methane boiloff gas to generate pressure, this makes a lot of sense, but not something who will work for an rocket but would be perfect for an moon lander or interplanetary injection burns.

[Nibb31]

Morpheus was a test platform for testing landing, throttling, and diversion techniques. The rocket part was pretty mundane, and not really the focus of the development project.

[Impaler]

This BDB nonsense is the stuff that know-nothing space cadets like to 'discover' and then think they know rocket science and can thumb their nose at NASA and justify why reality is not like science-fiction "It's some politicians/bureaucrats fault for canceling technology X". Nuclear rockets often serve the same role.

The BDB concept was for a PRESSURE FED engine, thouse things are abysmally low in ISP compared to tubo-pump engines which are NOT dumb. The BDB was never more then the fever dream of some engineer told to think outside the box, their is zero evidence that such a vehicle would deliver ANY payload to orbit let alone what it would cost. Remember getting to orbit is only just BARELY possible on Earth using chemical propulsion, if you 'dumb' down the rocket you sacrifice payload fraction, if you sacrifice it ALL it dose not matter how cheap the system is because it's doing nothing.

This is why SpaceX is on the right path, they are making the rocket higher performance and the manufacturing process as lean as possible so they have a good expendable rocket. But if they can get a high enough performance (the kind the Russians have had for decades) then they can sacrifice some of the performance to get the stages back and reuse them.

[Bill Phil]

Barely possible? You do realize that spacecraft have gone on trajectories that can get to the moon within hours on chemical propulsion.

It's funny how you praise the Russians, who have the closest thing to a BDB, but condemn BDBs themselves.

[Exoscientist]

You've gotta run the numbers to see if it works out. Lets say that your budding space program is building a rocket to land on the moon, and you have the option of using solid motors and hybrid solid-liquid ones, at 250 ISP, or Hydrogen, at 450 ISP.

18 km/s is what you need in delta-v. Let's assume that in each design, each stage is 5% non-fuel, and that the hydrogen rocket has 5 stages. The final one that returns being 5 tonnes dry.

Each stage has 3.6 km/s of delta-v.

So the rocket would weigh 430 tonnes at launch. Note that these numbers are optimistic, as in reality, the dry mass of the landing stage will be much higher. If the stages took 15% of the mass, we'd see a rocket at something like 900-1000 tonnes. Realistically, spent stages have quite significant masses, and low altitude performance is low.

Now if we have solids at 5% mass, we have 9 stages... And a whopping 15000 tonnes of booster. Remember, this is with the same formula that gave hydrogen boosters 430 tonnes for the same mission.

And 15% basically becomes madness. 53000 tonnes of MOAR BOOSTERS to accomplish the same task as 1000 of efficient rocket, and scaling solid motors to where the bottom stage weighs 30,000 tonnes wouldn't be simple or cheap to do.

but it's really not a good idea when modern hydrogen vehicles can get payload...

Thanks for the response. Keep in mind the saying "once you reach orbit, you're halfway to anywhere in the universe." The big problem with space exploration is just the huge expense of getting to orbit. You can use various efficient stages such as hydrogen-fueled ones once you reach orbit.

Bob Clark

- - - Updated - - -

This is nonsense. It barely hits 200m.

It's designed to manage the 2,000 m/s, or so, delta-v needed to make a lunar landing once delivered to lunar orbit, carrying a 500 kg payload. Calculate at the delta-v numbers you would get using smaller payloads. It's sufficient for suborbital, assuming you gave it a second engine so it can lift-off fully fueled in Earth's higher gravity or made it half size so the one engine will suffice.

Bob Clark

[Nibb31]

It's designed to manage the 2,000 m/s, or so, delta-v needed to make a lunar landing once delivered to lunar orbit, carrying a 500 kg payload. Calculate at the delta-v numbers you would get using smaller payloads. It's sufficient for suborbital, assuming you gave it a second engine so it can lift-off fully fueled in Earth's higher gravity or made it half size so the one engine will suffice.

Bob Clark

What exactly do you call suborbital? If I jump up and down, I'm technically doing a suborbital jump. Model rockets are suborbital.

If it's just to send a small science payload beyong the Karmann line, Von Braun could do it with an old V2, or you can use cheap solids, or even a small ASAT launched from an F-15.

Morpheus would be overengineered for that sort of thing because it was a *landing* research platform, not a *launch* research platform. The emphasis was on deep-throttling and ISRU-ready propellant and vacuum Isp. If you were to build a suborbital *launch* vehicle based on Morpheus' engine technology, then it wouldn't be Morpheus because it wouldn't need those things. You would basically and optimize it for atmospheric flight Isp, thrust, and easily-available propellant, which would make it a very different rocket.

Edited January 27, 2015 by Nibb31

[Northstar1989]

This BDB nonsense is the stuff that know-nothing space cadets like to 'discover' and then think they know rocket science and can thumb their nose at NASA and justify why reality is not like science-fiction "It's some politicians/bureaucrats fault for canceling technology X". Nuclear rockets often serve the same role.

The BDB concept was for a PRESSURE FED engine, thouse things are abysmally low in ISP compared to tubo-pump engines which are NOT dumb. The BDB was never more then the fever dream of some engineer told to think outside the box, their is zero evidence that such a vehicle would deliver ANY payload to orbit let alone what it would cost. Remember getting to orbit is only just BARELY possible on Earth using chemical propulsion, if you 'dumb' down the rocket you sacrifice payload fraction, if you sacrifice it ALL it dose not matter how cheap the system is because it's doing nothing.

Albert Einstain once said:

"If at first the idea is not absurd, then there is no hope for it."

It's good that you think the idea of a Big Dumb Booster is crazy, because that's *precisely* why it will work.

In all actuality, Big Bumb Boosters are far from unproven pipe-dreams. The Sea Dragon, the proposal that gave Big Dumb Boosters their name, was confirmed by NASA as being feasible and having accurate price-estimates:

"TRW conducted a program review and validated the design and its expected costs, apparently a surprise to NASA."

http://en.wikipedia.org/wiki/Sea_Dragon_%28rocket%29

The point of a Big Dumb Booster is that it relies on cheap technologies (like Pressure-Fed engines, which are used all the time in space programs- so don't call them absurd or terrible), and wide engineering margins (to bring down cost by allowing the components to be produced at much less specialized/precise factories), NOT that it has low ISP. The Sea Dragon still used LH2/LOX for the second stage- it just did so with a Pressure-Fed rocket engine (which admittedly hurt its ISP, but did not make it "abysmally low" by any means) and much wider engineering margins.

The bottom-line difference between a Big Dumb Booster and a "smart" booster is, once again, payload fraction and cost-per-kg. A Big Dumb Booster might have a 1% payload fraction instead of a 3% payload fraction, but makes up for it with less than 1/3rd the overall cost for a rocket of that size/mass. That's the basic principle- not low ISP, not low reliability (in fact, a Big Dumb Booster can be made MORE reliable than a "smart booster" thanks to much wider engineering margins). Of course, if you're also willing to sacrifice reliability, then you get something like the Aquarius (which also was a *small* rocket designed to take advantage of mass-production techniques...)

This is why SpaceX is on the right path, they are making the rocket higher performance and the manufacturing process as lean as possible so they have a good expendable rocket. But if they can get a high enough performance (the kind the Russians have had for decades) then they can sacrifice some of the performance to get the stages back and reuse them.

The Russian rockets are actually LOWER performance (as measured by mass and payload fractions) than American rockets. They are generally built to wider engineering margins, and are much closer to the Big Dumb Booster concept than anything utilized by NASA, the NSA, or Air Force- which is one of the reasons their rockets tend to get payload to orbit cheaper than American designs. It's ironic that you would mock the Big Dumb Booster concept, and then hold up the country that in real life comes closest to actually implementing it...

On the note of re-usability, yes, Space-X is going in the right direction for chemical rockets. But if you want to focus on performance and reusability, the ultimate end isn't chemical rocketry at all (which is more or less a peak technology- the only significant improvements we can make, like Full Flow Staged Combustion, tri-propellant fuel mixes, and Air Augmented booster stages; all yield relatively minor net improvements in performance and cost after the various trade-offs they require...) The ultimate end for high-performance rocketry is THERMAL rockets- specifically Microwave Thermal rockets and spaceplanes in the short run (Microwave Thermal has ISP's up to 850-1000 seconds when using Liquid Hydrogen, and 2-3 times the maximum TWR of chemical rockets), and Fusion Thermal spaceplanes somewhere decades or centuries down the road... (Fusion Thermal rockets have much better ISP than Microwave Thermal rockets, due to the much higher temperatures possible...)

And if you think Big Dumb Boosters are a pipe-dream, I can't imagine what you think of MICROWAVE THERMAL ROCKETS- even though the necessary technologies have all been demonstrated (one rocket scientist, Kevin Parker, literally built a test-thruster out of less than a hundred dollars of supplies in the lab- some of the parts literally being spare copper funnels from gardening... Gyrotrons are regularly used in some high-tech foundries and other parts of the metalworking industry, and a scientists demonstrated the ability to run a small toy electric helicopter off Microwave Power using a gyrotron and a rectenna in the 1960's...) Here's are some articles as an introduction to Microwave Beamed Power, by the way- I've actually done my research. I challenge you to read them

http://en.wikipedia.org/wiki/Beam-powered_propulsion

http://www.cnet.com/news/rocket-scientist-aims-to-relaunch-propulsion-technology/

https://physics.le.ac.uk/journals/index.php/pst/article/view/190

http://nextbigfuture.com/2014/02/escape-dynamics-and-microwave-power.html

http://authors.library.caltech.edu/3304/1/PARaipcp04b.pdf

http://authors.library.caltech.edu/3303/1/PARaipcp04a.pdf

NASA doesn't think the concept is dumb either, or they wouldn't have spent $2 million buying a 1-MW gyrotron unit *specifically* for studying Microwave Thermal propulsion:

http://callcenterinfo.tmcnet.com/news/2011/04/30/5477779.htm

Regards,

Northstar

[magnemoe]

Would not a solid stage fill this role just as well. They are even dumber than pressure feed engines?

For upper stages pressure feed might work as they are pretty small and don't need high powered engines on the other hand to keep the dry mass faction down and ISP up on them are critical.

[shynung]

Payload Assist Modules, which are generally Star-series small solid rockets, are occasionally used as upper stages already.

[Sampa]

After reading up on the Sea Dragon concept of the late 1960's, I've come to the conclusion that we're over-thinking this whole rocketry business.

http://en.wikipedia.org/wiki/Sea_Dragon_%28rocket%29

You see, the concept of Sea Dragon was simple- it's a LOT cheaper to build a large, heavy, powerful, sloppily-designed rocket that can lift the same payload to orbit with a lot greater inefficiency; than it is to build a carefully-engineered, narrow-tolerance rocket like most current US rocket designs. This concept is known, quite simply, as "the Big Dumb Booster principle".

http://en.wikipedia.org/wiki/Big_dumb_booster

So, why do we waste so much effort on ultra-engineered rocket designs, made out of fragile composites and high-tech alloys, when it would be a lot cheaper to slap something together out of corrugated steel and just launch the payload to orbit on THAT? It's not like we've been re-using any of our current rocket designs, anyways... The current US approach to rocketry strikes me a bit like building a Ferrari just so you can take it on one short drive off the edge of a cliff, when any old junker would do...

My best guess is that we're not ACTUALLY utilizing our current approach because it's cheaper or more effective (because it's not)- we're utilizing it because of lobbying (launch companies can charge more for high-performance designs) and because politicians want to generate more jobs in the space industry for political reasons...

Please discuss.

Regards,

Northstar

EDIT: See also the posts and links below on the Aquarius Rocket- a more modern cousin of the Sea Dragon proposed for initial launches in 2005/2006.

Simply put, Thrust to Weight Ratio or TWO. That is where the biggest problem with trying to get into space in the first place.

[Northstar1989]

Somedody make this a mod

I'm trying to get FreeThinker to do precisely that... We'll see if he bites...

- - - Updated - - -

Would not a solid stage fill this role just as well. They are even dumber than pressure feed engines?

For upper stages pressure feed might work as they are pretty small and don't need high powered engines on the other hand to keep the dry mass faction down and ISP up on them are critical.

The problem with Solid Rocket Motors is that their ISP is already so low that they don't really allow for much wider engineering margins and less precise construction without losing their usefulness (beyond as strap-on boosters) altogether. A low-precision, high engineering-margin Kero/LOX Big Dumb Booster stage still has *MUCH* better ISP than a high-precision Solid Rocket Motor (and is cheaper, since the manufacturing can be much less precise).

Not to mention, Solid Rocket Fuel isn't actually cheap- what makes SRB's cheap is that no work needs to go into designing separate engines or turbopumps or such- it's just "light and pray". The casing, control, and thurst-gimballing components still need to be built to *VERY* precise engineering margins in order to get much useful Delta-V out of them when using them as anything but strap-on-boosters...

In Summary: A SRB built to Big Dumb Booster standards would have very low ISP *AND* terrible mass-fraction. Not exactly a recipe for lots of Delta-V. And you would see less in the way of cost-savings than with a Big Dumb Booster liquid stage, as less of the cost of a SRB is invested in the dry mass components anyways- and more of it is invested in the relatively much more expensive fuel...

SRB's built to low precision and engineering standards could have exactly *ONE* use, and one use only- as strap-on boosters for the launch stage (where they are adding to Delta-V as long as their TWR exceeds the rest of the rocket...) And mant strap-on SRB's actually already are built to Big Dumb Booster standards anyways- many use corrugated steel with wide engineering margins for their casings due to its cheapness to machine (even though it's much heavier than aluminum)- and thus ALREADY qualify as Big Dumb Boosters... But due to the cost of the propellant itself, an SRB built like a Big Dumb Booster will never be as cheap as a liquid rocket built like a Big Dumb Booster...

Regards,

Northstar

P.S. Those of you suggesting that Space-X style reusables will be more cost-effective than Big Dumb Boosters for the launch stage might very well be right- especially if Space-X succeeds in developing Full Flow Staged Combustion engines- but what about the upper stages? Space-X may have *plans* to re-use their upper stages, but I *highly* doubt this is going to turn out to be nearly as economical as they hope...

Don't get me wrong- I'm sure they can do it (in fact, I regularly build Space-X style launch systems with reusable launch *and* upper stages with Real Solar System 64K, FAR, and Deadly Re-Entry installed all the time, and I'm not a professional engineer... Admittedly, my upper stages usually lack landing legs, and often make a soft-landing but then lose parts after falling over...) But, I suspect the cost of replacing all the thermal tiling and such between launches will be much higher than just building a cheap (Big Dumb Booster cheap, NOT high-precision engineering cheap) expendable upper stage and de-orbiting it (or boosting it to a graveyard orbit) after the payload is in place... Either way you're going to need to lift extra mass to orbit (either for having a Big Dumb Booster upper stage, or for all the heat-shielding and control systems necessary for a recoverable upper stage), but your cost-savings from cheap construction/manufacturing may be greater than your cost-savings from re-usability, when it comes to upper stages...

Edited January 29, 2015 by Northstar1989

[78stonewobble]

How big, how dumb and how many?

Well, my personal guess is as big as you can reasonably make an engine and rocket body and ensure a reasonable quality, while still being able to mass produce (really churn them out like liberty ships).

[The Yellow Dart]

P.S. Those of you suggesting that Space-X style reusables will be more cost-effective than Big Dumb Boosters for the launch stage might very well be right- especially if Space-X succeeds in developing Full Flow Staged Combustion engines- but what about the upper stages? Space-X may have *plans* to re-use their upper stages, but I *highly* doubt this is going to turn out to be nearly as economical as they hope...

I don't think big dumb boosters could ever beat out even just 1st stage recovery at this point. If you take a Falcon 9, that thing has 9 Merlin rocket engines on the 1st stage vs. 1 Merlin-Vac on the second. Recovering 9 out of 10 engines plus the much larger of the 2 tank sections will surely beat out big dumb boosters no matter how cheap you are able to make it. At the end of the day, throwing away a cheaper engine is only slightly less dumb than throwing away an expensive one. And Merlins are *theoretically* capable of 40 launch cycles before any significant refurbishment is necessary other than shuttling back to the launch site and refueling. Even if it turns out to be just half that many uses, you are still getting 20 times the use out of a single booster. Can we really decrease the cost of building an engine by a factor of 20? Seems doubtful.

Ideally, I think it would have been smart for NASA to pick up the big dumb boosters idea as an interim in the 70s and 80s while true reusability was in development, but of course, ideally, the example of true reusability we were given would not have been the space shuttle.

[Rakaydos]

40 firing cycls, not 40 launches. The center Merlin fires 4 times per launch.

Still, if they rotate engines every launch, even a falcon heavy will only need togo through (8+4) x3 firings before being rotated to the upper stage... which by my matth is exactly 40 firing cycles befor

E discarding.

[The Yellow Dart]

40 firing cycls, not 40 launches. The center Merlin fires 4 times per launch.

Still, if they rotate engines every launch, even a falcon heavy will only need togo through (8+4) x3 firings before being rotated to the upper stage... which by my matth is exactly 40 firing cycles befor

E discarding.

That's true, I forgot about the landing firings. But according to Musk's AMA the other day, 40 cycles isn't a throw-away limit but the point where certain components will need replaced. He made it sound like they hadn't yet discovered the maximum life of the engine. How much one of those refurbishments will cost compared to the outright cost of a new engine is a good question, maybe for the next time he does another AMA.

[Impaler]

Northstar1989 : An airospace cost estimate before something is done is hardly worth anything, they are routinely off by orders of magnitude. And frankly NASA would be completely negligent to accept someone else cost estimate for ANYTHING. From what I've been able to find these costs estimates assumed absurd launch rates of nearly a million tons a year and partial reuse. These seem to be the studies.

http://neverworld.net/truax/Sea_Dragon_Concept_Volume_1.pdf

http://neverworld.net/truax/Sea_Dragon_Concept_Volume_3.pdf

Stage propellent fractions are 88% on both the first and second stage, that's terrible. ISP claims are 242 for the first stage using Kero/LoX 409 for second stage Hydro/Lox, both of which are utterly absurd at the time, they are basically saying their pressure fed engine first stage engine would be just 18 sec lower ISP then the F-1 on the Saturn V and the second stage engine would be only 9 seconds less then the J-2. And all of this is done with chamber pressures that are an order of magnitude lower then what a pump-fed engine produces. I don't think these ISP numbers for pressure-fed were remotely achievable in that day and are probably not achievable now, this is why every nation has abandoned pressure-fed engines and move to pump-fed when they go beyond small sounding rockets.

Your point about Russian rockets is dead wrong. Russian KeroLox rocketry has THE HIGHEST ISP performance in the world among Hydrocarbon rockets that is why we are BUYING THEM. An American made HydroLox rocket dose achieve a higher ISP because of it's fuels molecular weight but at enormous costs. The Russians and SpaceX rightly decided that it is not worth the cost.

Russian/SpaceX style rockets are absolutely NOT big Dumb Boosters, anyone trying to claim their success as somehow retroactively validating SeaDragon has their head up their ass. These rockets are not high tolerance low performance 'dumb' things, they are LOW tolerance HIGH performance, they are made in large NUMBERS to bring manufacturing costs down. This is exactly the opposite approach from what Sea Dragon would have employed with one huge engine.

As for your throwing out bunch of other unrelated propulsion concepts and challenging me that I must some how be opposed to them as well simply because they are 'advanced' is typical space-cadet thinking, aka anyone who disagrees with you is a defeatist and must think that no technology is worth pursuing.

SeaDragon is just a bad DESIGN, the things it claims will save costs are not remotely likely to do so, like operating at Sea, anyone who knows anything about logistics can tell you that for for the same size logistical task the sea is many times more expensive. In SeaDragon they are just UNABLE to do it on land and then claim the sea is 'cheap', possible things are always going to look cheap compared to impossible things. The question is if the launch costs of SeaDragon at Sea would be cheaper then a non-dumb rocket on land, considering that SeaDragon was going to need a Nuclear Aircraft Carrier to make it's fuel and oxidizer onsite I can hardly call that cheap. The only pieces of infrastructure your replacing in the ocean area he concrete pad and the vehicle crawler, everything else that's part of a launch complex now has to be done from logistical ships at 10x the costs.

Edited January 30, 2015 by Impaler

[Rakaydos]

That's true, I forgot about the landing firings. But according to Musk's AMA the other day, 40 cycles isn't a throw-away limit but the point where certain components will need replaced. He made it sound like they hadn't yet discovered the maximum life of the engine. How much one of those refurbishments will cost compared to the outright cost of a new engine is a good question, maybe for the next time he does another AMA.

Actually, I was saying that every launch there is 40 separate engine firings, and one engine simply isnt recovered. By rotating the more-worn engines to the unrecoverable stage, they avoid the cost of refurbisment while getting the most out of their design.

[Northstar1989]

Your point about Russian rockets is dead wrong. Russian KeroLox rocketry has THE HIGHEST ISP performance in the world among Hydrocarbon rockets that is why we are BUYING THEM. An American made HydroLox rocket dose achieve a higher ISP because of it's fuels molecular weight but at enormous costs. The Russians and SpaceX rightly decided that it is not worth the cost.

Russian/SpaceX style rockets are absolutely NOT big Dumb Boosters, anyone trying to claim their success as somehow retroactively validating SeaDragon has their head up their ass. These rockets are not high tolerance low performance 'dumb' things, they are LOW tolerance HIGH performance, they are made in large NUMBERS to bring manufacturing costs down. This is exactly the opposite approach from what Sea Dragon would have employed with one huge engine.

Impaler, your posts are consistently rude and offensive, and totally off-base... You also mis-represent my posts. For instance, I said:

The Russian rockets are actually LOWER performance (as measured by mass and payload fractions) than American rockets.

Note the words *as measured by mass and payload fractions*.

ISP is entirely separate than mass/payload fractions. I am well-aware that Russian rocket engines have the best hydrocarbon-ISP in the world (thanks mainly to more advanced metallurgy, and higher chamber pressures). That is *NOT* the same thing as having good mass-fractions. In fact, one way to IMPROVE specific impulse, while HURTING mass-fraction is to make use of a higher-chamber pressure in your rocket engines. This in turn requires thicker engine walls, which requires an engine to have higher mass. You may get better stage Delta-V (up to a certain point) due to better ISP, but you actually hurt the stage mass-fraction...

Hard Numbers:

Saturn V payload-fraction: 4.6%

Soyuz-FG payload-fraction: 2.3%

The Soyuz-FG has a launchpad-mass of 305 metric tons (305,000 kg). Its payload capacity is 7.1 metric tons (7100 kg), meaning its payload-fraction is *just* above 2.3%. The Soyuz-FG made its first flight in 2001, whereas the Saturn V had its maiden flight in 1967. The Soyuz-FG has the advantage of 24+ years of technological progress on its side...

Regards,

Northstar

Edited February 2, 2015 by Northstar1989

[Kryten]

The Soyuz-FG is a Soyuz-U with flight computer upgrades, it's had no propulsion system upgrades since some time in the 60s. You also ignore the fact that larger rockets are inherently more efficient, something you do clearly know about as it is one of the basis' for this thread.

[Northstar1989]

Stage propellent fractions are 88% on both the first and second stage, that's terrible.

Low propellant-fractions is the WHOLE POINT of a Big-Dumb Booster. You save more in manufacturing-costs with the wider engineering margins than you incur in costs from needing a larger total rocket. The very fact that the mass-fractions were bad makes the rocker MORE feasible (they aren't trying to push the limits of technology, NOT less... The fact that you are criticizing this shows you haven't been readings my posts at all, just behaving rudely and offensively, and in a manner that should not be permissible on the KSP Forums. I suggest you take a little more care to be respectful of your fellow KSP-er's...

ISP claims are 242 for the first stage using Kero/LoX 409 for second stage Hydro/Lox, both of which are utterly absurd at the time, they are basically saying their pressure fed engine first stage engine would be just 18 sec lower ISP then the F-1 on the Saturn V and the second stage engine would be only 9 seconds less then the J-2.

Your numbers are wrong (the F-1 has a sea level ISP of 263 seconds, not 262 seconds; J-2 had an ISP for 421 seconds, not 418 seconds, so ISP would be 19 and 12 seconds lower- small differences like this mean a LOT in rocketry), and you are skewing the numbers to try and make your point seem more valid...

And all of this is done with chamber pressures that are an order of magnitude lower then what a pump-fed engine produces. I don't think these ISP numbers for pressure-fed were remotely achievable in that day and are probably not achievable now, this is why every nation has abandoned pressure-fed engines and move to pump-fed when they go beyond small sounding rockets.

The chamber pressure for the Sea Dragon's first-stage engine would be 300 psi, compared to 1015 psi for the F-1 engine. That is *NOT* an "order-of-magnitude" difference (generally defined as a 1:10 ratio or more), the pressure difference is only 1:3.

Not only that, but 242 seconds is an entirely reasonable ISP for a chamber pressure of 300 psi, *especially* given that the Sea Dragon had an 8-year technology advantage over the F-1 (the F-1 was proposed and designed in 1955, whereas the Sea Dragon first-stage engine was proposed in 1963...) and relied on a much more expandable nozzle (expands from 7:1 to 27:1 as the engine climbs, compared to 10:1 to 16:1) which allowed for better performance at high-altitude. The formula for exhaust-velocity, by the way:

Ve = SQRT[(2*k/(k-1))*(R'*Tc/M)*(1-(Pe/Pc)^(k-1)/k)]

Or, alternatively, a simplified version:

[1 - (pe/pc)^(k-1)/k] / (k-1)

Notice the term "Pe/Pc"? This is the *ONLY* term in the entire equation where chamber-pressure comes into play- and even then not directly. What determines exhaust-velocity (and thus Specific Impulse) is not the chamber-pressure, it is the ratio between the chamber pressure and the exhaust pressure. This is determined by nozzle-ratio.

The Sea Dragon design had an initial nozzle-ratio of 7:1, with an expandable nozzle that could expand up to 26:1 in the upper atmosphere. The Rocketdyne F-1 engine (the launch stage engine of the Saturn V) had a nozzle-ratio of 10:1, with an expandable nozzle that could expand up to 16:1 in the upper atmosphere. In neither case is the exhaust-stream ever over-expanded, but the Rocketdyne F-1 exhaust stream loses a lot of potential I_sp in the upper atmosphere by having a lower nozzle-ratio.

The Sea Dragon was TSTO (Two Stage To Orbit), so a higher maximum nozzle-ratio was designed in despite the extra weight as the launch stage would be flying higher into the atmosphere than the 3-stage Saturn V... The larger expandable nozzle; a "drag skirt" specifically designed to reduce the terminal velocity of the Sea Dragon's launch stage to under its maximum designed crash-tolerance of 300 ft/s during re-entry, such as to allow first-stage recovery with a purely drag-based landing in the ocean; and several design features that allowed the Sea Dragon to survive immersion and erection (being stood on end prior to sea-launch) in the ocean without ill-effects, all contributed to the "low" mass-fraction of 88%, as did the much wider engineering-margins and *MUCH* sturdier design (corrugated-steel, which as already stated, allowed the rocket to survive splashdowns at up to 300 ft/s without damage to any of the rocket's components) to allow for cheaper construction...

All other terms in the equation besides nozzle-ratio cancel out when comparing the Sea Dragon and Saturn V except for T_c (chamber temperature), as they are based on the fuel used, which is the same (RP-1/LOX) in both cases.

Thus, the only factors that differ which affect I_sp are the nozzle area-ratio (which is 7:1 instead of 10:1 for the Sea Dragon main engine vs. the Rocketdyne F-1 engine...) and the chamber temperature.

Regards,

Northstar

- - - Updated - - -

The Soyuz-FG is a Soyuz-U with flight computer upgrades, it's had no propulsion system upgrades since some time in the 60s. You also ignore the fact that larger rockets are inherently more efficient, something you do clearly know about as it is one of the basis' for this thread.

With all due respect, you're wrong:

Work on the 14D21 and 14D22 engines started in 1986, with a preliminary design completed in 1993.

The RD-107A and RD-108A are improved versions of the original RD-107 engine (originally designed in 1954-1957) that were designed from 1986 to 1993. They:

incorporate a new injector head design to increase specific impulse.

Theses engines were first utilized in May 2001 to launch a Progress cargo spacecraft to the ISS, according to Wikipedia. Hardly sounds like a 1950's engine to me...

Regards,

Northstar

Edited February 2, 2015 by Northstar1989

[architeuthis]

Notice the term "Pe/Pc"? This is the *ONLY* term in the entire equation where chamber-pressure comes into play- and even then not directly. What determines exhaust-velocity (and thus Specific Impulse) is not the chamber-pressure, it is the ratio between the chamber pressure and the exhaust pressure. This is determined by nozzle-ratio.

I agree with your general point, but chamber pressure is important in that it is used to determine combustion curves for different adiabatic flame temperatures vs. mixture ratios. In some cases the effect of chamber pressure on engine performance is extremely significant. In other words, optimal values of k, m, R' and T all depend on choice of Pc!

Here is a link to some slides I made on the subject, if you're interested!

Edited February 3, 2015 by architeuthis

[magnemoe]

The Sea Dragon was TSTO (Two Stage To Orbit), so a higher maximum nozzle-ratio was designed in despite the extra weight as the launch stage would be flying higher into the atmosphere than the 3-stage Saturn V... The larger expandable nozzle; a "drag skirt" specifically designed to reduce the terminal velocity of the Sea Dragon's launch stage to under its maximum designed crash-tolerance of 300 ft/s during re-entry, such as to allow first-stage recovery with a purely drag-based landing in the ocean; and several design features that allowed the Sea Dragon to survive immersion and erection (being stood on end prior to sea-launch) in the ocean without ill-effects, all contributed to the "low" mass-fraction of 88%, as did the much wider engineering-margins and *MUCH* sturdier design (corrugated-steel, which as already stated, allowed the rocket to survive splashdowns at up to 300 ft/s without damage to any of the rocket's components) to allow for cheaper construction...

Regards,

Northstar

As its build like a ship it would be resistant to seawater. However a ship would not survice an 360 km/h impact with the sea, record here is 80 km/h and its an small lifeboat designed to dive down to brake. If engine was restartable it should have been posible to do a braking burn to reduce speed to manageable levels before splashdown, it should be able to handle a pretty rough landing but its limits.

However reuseablity argue against the cheap booster idea, you still want it robust and easy to maintain and would be willing to send more money on this, you would also spend money on increasing performance.

base issue is that huge rockets are effective as many of the costs are fixed and the inverse square law helps you. Downside is trying to get cargo for enough sea dragon launches.

[Northstar1989]

As its build like a ship it would be resistant to seawater. However a ship would not survice an 360 km/h impact with the sea, record here is 80 km/h and its an small lifeboat designed to dive down to brake. If engine was restartable it should have been posible to do a braking burn to reduce speed to manageable levels before splashdown, it should be able to handle a pretty rough landing but its limits.

However reuseablity argue against the cheap booster idea, you still want it robust and easy to maintain and would be willing to send more money on this, you would also spend money on increasing performance.

base issue is that huge rockets are effective as many of the costs are fixed and the inverse square law helps you. Downside is trying to get cargo for enough sea dragon launches.

The maximum survivable splashdown speed of the Sea Dragon was 300 ft/s- which equates to about 329 km/h, not 360. However, I agree that it's a rather impressive number. The thing to remember about a rocket is that it's already built to survive substantial g-forces anyways: just normally rockets are very light and thus the total forces survived aren't that great in magnitude (for analogy: a needle can survive far more g's than a large wooden block, but the block can survive greater *magnitudes* of force...) However the Sea Dragon was *VERY* sturdily built, and unlike a boat, has very little mass besides the structural elements (which were quite heavy for a rocket) once the fuel is burnt-out. I have little doubt that a hollow metal tube the size of a battleship could survive a 328 km/h splashdown (that's a little over 91 m/s, for reference) if it didn't have *ANY* cargo or fuel mass weighing it down...)

There *were* provisions for minor repairs to the structure after splashdowns (the Sea Dragon was designed for easy repair), however, and they did design a drag-skirt to reduce splashdown speed to something like 20 or 30 m/s- so they clearly weren't expecting to push the maximum structural tolerances on a regular basis...

As for cargo-mass, the Sea Dragon would have been *quite* useful as a fuel-tanker or for launching large fuel-depots to LEO. The payload-capacity of the Sea Dragon was less than twice that of the Saturn V, despite being more than four times as massive on the launchpad: and the vast-majority of mass necessary for missions beyond LEO is fuel- so it's not like you couldn't have burned through that fuel mass in a fairly small number of missions (*especially* if you were using lower-ISP storable propellants like Kerosene, instead of cryogenics, and launching mission vehicles with only the necessary LOX onboard...) It also could have launched the ISS in just a handful of launches, for a *fraction* of the total cost of the Shuttle (or even using Saturn V rockets to launch the ISS).

Regards,

Northstar

Edited February 4, 2015 by Northstar1989

[Bill Phil]

Just had a thought:

What about BDSBs? Big Dumb Solid Boosters? Many all solid rockets have been flown, like the Scout and Minotaur rockets, as well as the Athena family. So why not try to make BDBs with solids? It could have a payload of one tonne and a total mass of over 100 tonnes with more than two stages...

Edited September 30, 2015 by Bill Phil

[Kryten]

Just had a thought:

What about BDSBs? Big Dumb Solid Boosters? Many all solid rockets have been flown, like the Scout and Minotaur rockets, as well as the Athena family. So why not try to make BDBs with solids? It could have a payload of one tonne and a total mass of over 100 tonnes with more than two stages...

How does that have anything to do with the BDB concept? It's not even Big, that's basically just Vega, CZ-11 or Minotaur-IV.