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  #1  
Old 12-05-2009, 03:07 PM
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CPMcGraw CPMcGraw is offline
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While browsing around the 'net, I came across the technical specs on the Russian RD-191 engine assembly. Unless I've mis-read the power numbers, this cluster produces 1.7 M pounds of thrust, and is capable of throttling down to 30% of its rating.

We all know how cool the Saturn 1B models look with a four-engine cluster...

The original 1B required eight engines to produce 1.6 M pounds of thrust...

So, the idea would be to swap the RD-191 into the basic airframe of a Sat1B. I'm thinking there would be a good weight reduction in the mounting hardware for the RD-191, as well as a reduction in the plumbing requirements. Also, might there be other weight reductions in modern alloys and welding technologies?

Additionally, what about the S-IV-B stage? Could similar weight reductions and power improvements be seen using something like the RD-180 (Atlas V engine) in place of the J-2?

Would such a modernized Saturn be enough to lift the Orion into LEO?

Foo for thought...
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  #2  
Old 12-05-2009, 11:08 PM
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You should post this over on http://forum.nasaspaceflight.com/index.php?board=37.0 and you'd get an answer from REAL rocket scientists who do these sort of calculations for breakfast...

From the sound of it, yes. It's problems would be more political than technical... and assuming the RD-191 was licensed for US manufacture. Toolup costs would be significant, but it would be cheaper than designing and building an entirely new engine.

Personally I think kerosene is the way to go. Solids are becoming more and more of an environmental concern because of perchlorates in the water and stuff down there, and they are FAR more useful for pure raw brute power for heaving enormous payloads off the pad-- for a manned vehicle first stage I think that decision will be one we'll someday regret-- probably at some astronaut funerals... Kerosene has the density on it's side, meaning smaller and lighter structures than a low-density hydrogen fuelled first stage, and liquid engines are throttleable and can be shut down in flight at any time, making aborts much safer. They are also FAR more benign in their failure modes than solids are, and typically much easier to detect problems requiring an abort before things start to come apart at high speed.

There were concepts vetted to replace the "cluster's last stand" amalgam of tanks Saturn I/IB first stage with a new unitary fuel tank and oxygen tank of 22 foot diameter mated with an intertank to form a new first stage for Saturn I. It would have been MUCH more weight-efficient and cheaper to produce, but would have had significant design, development, testing, and certification costs compared to the existing Saturn I clustered first stage. There were also plans to power it with a single or pair of F-1 or F-1A engines. Such a first stage would have made a TERRIFIC liquid strap-on booster for Saturn V as well, using a pair of them... imagine-- NINE F-1 ENGINES FIRING AT LIFTOFF!!!! ISS to orbit in a single launch! (or darn close to it!)

Using a kerosene upperstage is certainly doable, though the performance is less due to the lower ISP of the kerosene engines. Hydrogen is REALLY what you want for upperstages because you can squeeze a LOT more performance out it... and hydrogen would be necessary for a HLV EDS stage to leave orbit (I suppose that theoretically, a kerosene EDS MIGHT be possible, but would have pretty crummy performance) IF you need hydrogen infrastructure as part of your GSE anyway, better to go with a hydrogen upperstage as well and use the common upperstage engines on your EDS as well. Either way, kerosene or hydrogen, you're talking a whole new second stage for your CLV...

Personally, I LIKE the kerosene proposal NASA has recently vetted against Ares V-- though they have NO intention of building it... it's strictly a smokescreen in their parlor game to get Ares V funded no matter what... but actually, IF I were king, I'd go one of two ways--

Either immediately switch to dual launch using a DIRECT style SDLV using as much of the existing shuttle stack as possible with a new upperstage using off the shelf RL-10 engines, just as Saturn I's second stage, the S-IV did...

OR, I'd say "go clean sheet" and develop a kerosene booster much like what you're suggesting... one that could be used as a 'single stick' with an upperstage to launch your crew, and strap 3 of them together, possibly with a larger twin-engine core if necessary, but using the crew launcher first stage as strap on boosters, to launch (hopefully) a second stage virtually identical to the second stage of the crew launcher, just equipped as an EDS... THAT would make the most sense to me, and get us off the 'SRB addiction' and yet still have the synergies of commonality between the crew launcher and cargo launch rockets...

But that makes too much sense... NASA wants to do the most difficult and expensive thing possible that makes the least sense and takes the most time and money to do... because that's how you make a jobs program that lasts 20 years... The rocket is just a means to an end-- it's the JOBS program that counts... all that money spent on rockets is SPENT ON THE GROUND...

OH well... later! OL JR
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Old 12-06-2009, 04:56 AM
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I think the Russians had an excellent idea with their all-liquid propellant Energia/Buran ULS (Universal Launch System). The LOX/LH2 core stage was boosted by four LOX/kerosene boosters (Zenit first stages) that had provisions for parachute recovery (not tested during the program's lifetime). The ULS could carry either a large payload and its orbital insertion stage *or* a Buran orbiter which had its own orbital insertion/maneuvering engines.

When carrying the orbiter, the ULS was capable of multiple abort options (which were safer than those available for the US Space Shuttle because the ULS's LOX/kerosene boosters could be shut down), including Return To Launch Site (RTLS), Abort Once Around (AOA), Abort To Orbit (ATO), and the Russian equivalent of the US Space Shuttle's Trans-Atlantic Abort Landing (TAL), in which the orbiter would follow a suborbital ballistic trajectory to a downrange landing site.

There has been some talk in recent years about reviving the ULS for supporting joint Russian/US manned lunar, asteroid, or Mars missions, but unfortunately it will probably remain just that--talk.
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Last edited by blackshire : 12-06-2009 at 04:58 AM. Reason: This ol' hoss done forgot somethin'.
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Old 12-06-2009, 08:50 AM
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Quote:
Originally Posted by CPMcGraw
So, the idea would be to swap the RD-191 into the basic airframe of a Sat1B. I'm thinking there would be a good weight reduction in the mounting hardware for the RD-191, as well as a reduction in the plumbing requirements. Also, might there be other weight reductions in modern alloys and welding technologies?


Craig, what you're proposing could certainly be done. I suspect you'd be better off starting from a clean sheet of paper and just designing the launcher you want for the mission you're trying to fly, though. Launch vehicles are designed to fly specific missions. The loads and environments that derive from propulsion and flight profiles are used to design structure, set the bounds for certification tests, etc. Margins are usually trimmed to maximize performance. Make any significant change and you typically have to change everything else ... in effect, you're designing a new vehicle.
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Old 12-06-2009, 09:04 AM
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Originally Posted by luke strawwalker
IF you need hydrogen infrastructure as part of your GSE anyway, better to go with a hydrogen upperstage as well and use the common upperstage engines on your EDS as well.


In fact, since hydrogen offers higher ISP than kerosene and hydrogen infrastructure already exists while kerosene does not, you'd be better off using LOX/LH2 for the booster as well and maximize performance all around. You'd have to trade the long-term costs associated with the propellants themselves against that of installing and maintaining parallel H2/K infrastructure, living with reduced performance (lower mass to orbit drives more launches), etc to know if it's the best approach from an economic perspective.
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Old 12-07-2009, 12:04 AM
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Originally Posted by wilsotr
In fact, since hydrogen offers higher ISP than kerosene and hydrogen infrastructure already exists while kerosene does not, you'd be better off using LOX/LH2 for the booster as well and maximize performance all around. You'd have to trade the long-term costs associated with the propellants themselves against that of installing and maintaining parallel H2/K infrastructure, living with reduced performance (lower mass to orbit drives more launches), etc to know if it's the best approach from an economic perspective.

Hydrogen-fueled first stages have three disadvantages:

[1] While LH2 does have a higher ISP than RP-1 (rocket-grade kerosene), its ISP advantage (due to its higher exhaust velocity) can only be realized in near-vacuum and vacuum conditions. At sea level and low altitudes, the low-molecular mass exhaust of a hydrogen/oxygen rocket engine (superhot steam and some unburned hydrogen) translates into lower liftoff thrust for the same engine size. Lower liftoff thrust in turn means that less rocket airframe mass and less payload mass can be lifted.

[2] The requirement for larger LH2/LOX first stage engines (for the same liftoff thrust as RP-1/LOX engines) also cuts into the vehicle's payload mass. The size difference between hydrogen and kerosene engines of the same thrust level is significant. In the 1960s Aerojet designed the M-1, a hydrogen/oxygen engine (see: http://en.wikipedia.org/wiki/M-1_%28rocket_engine%29 and http://www.secretprojects.co.uk/for...hp?topic=8193.0 ) whose baseline design was rated at 1.5 million pounds of thrust (the same as each one of the Saturn V's five F-1 kerosene/oxygen first stage engines). The M-1 was at least twice the size of the F-1, which would make it about 8 times as heavy. While it was intended primarily for upper stage use, Aerojet also studied first stage versions of the M-1 that would have operated at a much lower ISP at sea level.

[3] Due to liquid hydrogen's low density, its fuel tank must be very large, which increases its mass. Since a multi-stage rocket's first stage has the largest and heaviest fuel tank to be found in the vehicle, a hydrogen-fueled first stage is bigger and has a larger mass when empty (and thus a poorer mass fraction) than a kerosene-fueled first stage of comparable performance. Such a larger first stage also encounters greater aerodynamic drag in the lower atmosphere. These factors, combined with LH2/LOX engines' lower sea-level thrust, larger size, and higher mass (for the same thrust as RP-1/LOX engines) makes LH2/LOX a less attractive propellant combination for first stages.

In a two-stage satellite launch vehicle, the first stage provides the brawn to lift the vehicle away from the Earth's surface, boost it through the denser regions of the atmosphere, and get the vehicle on a more nearly horizontal trajectory. Once it has done its job and separated, the now much lighter second stage sprints through the near-vacuum of the upper atmosphere and then the vacuum of space, using its higher exhaust velocity to accelerate itself and its payload to orbital velocity.

It is no coincidence that big, heavy-lift launch vehicles like the Saturn V and Proton were/are often called "workhorse vehicles." Like a slow but immensely strong draft horse, their powerful but lower-ISP first stages were designed to overcome gravity by sheer brute force and lift their upper stages to altitudes and velocities from where they could complete the harness race-like climb to orbit (in which, as with a harness horse race, the greatest acceleration occurs at the finish) with less expenditure of energy each second (although upper stages often contain more *total energy* in their propellants than first stages).

To summarize, RP-1/LOX is a superior propellant combination for high-thrust "muscular" first stages that must overcome gravity and aerodynamic drag, while LH2/LOX provides a longer and more efficient (although lower-thrust) push in the near-vaccum of the upper atmosphere and in the vacuum of space.
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Last edited by blackshire : 12-07-2009 at 12:36 AM. Reason: This ol' hoss done forgot somethin'.
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Old 12-09-2009, 12:48 PM
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Blackshire I'm not sure I understood your first point, but in general you are correct: the higher Isp offered by LH2 is offset, at least in part, by the additional vehicle mass required for high-volume LH2 tankage, cryo insulation, and etc. Drag eats into the Isp advantage too if the vehicle core diameter is increased to grow tank volume - this unless payload requirements drive a large-diameter shroud of some kind anyway. Everything else being equal, a short-burn, high-thrust RP-1 first stage coupled with a high-Isp LH2 second stage (SSME class) would be the theoretically optimum way to maximize mass to orbit.

RP-1 infrastructure doesn't exist at LC-39 though, so it would have to be designed and installed. That drives up initial system procurement costs. Then it has to be maintained in parallel with the LH2 infrastructure ... that has implications for long-term recurring costs. If the RP-1 design has a high-enough initial thrust to eliminate the solids there's a savings buried there to offset those, but that's all part of the trade. There is no domestically-produced high-thrust RP-1 engine available. You can continue with the licensing / foreign-produced approach taken for the RD-180, of course, but really need to think about the long-term implications of that for this vehicle and the US industrial base and weigh those against the development cost associated with producing something new.

As it turns out, the Isp of SSME's is high enough that a long-burn SSME-core coupled with 5-segment solids produces a vehicle with about the same mass-to-orbit capability as the RP-1 liquid-only vehicle - this without the RP-1 infrastructure and engine development (or licensing) risks, but a whole different set of design issues, procurement and recurring costs. The choice of one or the other as "optimum" isn't very clear until you've done the trades and factored in all those other considerations. Even then it isn't very clear.
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Old 12-09-2009, 01:23 PM
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Sigh. Where is von Braun when you need him?

My preference is an RP-1/LOX first stage, with second stages and beyond being LH2/LOX (hmmmm..... sounds like the good ol' Saturn V). A lot of this is political, perhaps nearly all. The space program should be considered a national capital asset, and part of our national identity. We can either be a leader or sit on the side lines.

My reference to WvB means that not only was he a excellent engineer (and engineering manager), he was also a charismatic champion for the cause: America in Space is part of our manifest destiny. If People Magazine had been around in the '50s, he would have probably been on the cover of it a time or two. I look at the popular landscape today and there is no one like that today.

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Old 12-09-2009, 08:12 PM
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Originally Posted by GregGleason
The space program should be considered a national capital asset, and part of our national identity.


Totally agree ......
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Old 12-09-2009, 09:24 PM
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Quote:
Originally Posted by GregGleason
Sigh. Where is von Braun when you need him?

My preference is an RP-1/LOX first stage, with second stages and beyond being LH2/LOX (hmmmm..... sounds like the good ol' Saturn V). A lot of this is political, perhaps nearly all. The space program should be considered a national capital asset, and part of our national identity. We can either be a leader or sit on the side lines.

My reference to WvB means that not only was he a excellent engineer (and engineering manager), he was also a charismatic champion for the cause: America in Space is part of our manifest destiny. If People Magazine had been around in the '50s, he would have probably been on the cover of it a time or two. I look at the popular landscape today and there is no one like that today.

Greg


What's amazing to me, in having read and followed a lot of the design issues and tradeoffs over the last couple years or so on nasaspaceflight.com, is just how good Von Braun and his team actually were... they REALLY hit the sweet spot on so many levels... the Saturn V was REALLY a great design... The RP-1 first stage had AWESOME performance and the choice of RP-1 gave it the smallest size possible for a slightly reduced ISP by using kerosene, but which was a good tradeoff considering the superior thrust characteristics blackshire noted above from liftoff until exiting the lower atmosphere. It then staged at just about the perfect point, getting rid of the massive first stage and igniting the high-performance hydrogen fuelled common bulkhead, exceptionally light for it's size second stage to take the vehicle almost to orbit, before staging to the restartable third stage which rapidly completed the ascent before shutting down and later restarting to push the stack through TLI, minimizing the 'dead weight' which directly reduces lunar performance pound for pound...

Ares V uses relatively low ISP (for hydrogen engines) RS-68's, which requires them to carry more fuel for a given job, which reduces payload capacity (the weight of extra fuel that could have been devoted to payload if the engine was more efficient, but such engine would cost more). The use of hydrogen on the core means the core stage has to be absolutely ENORMOUS, when coupled with the thirsty sorta low ISP RS-68's and the fact that it's basically performing part of the first stage's job AND the second stage on the Saturn V, while still carrying all that dead weight of tankage that SHOULD have been disposed of in a staging event, but isn't, because Ares V is two stage instead of three stage. Carrying all that dead weight along further into the flight cuts into performance significantly. Using poor ISP SRB's as a 'first stage' of sorts while burning the thirsty sorta-low ISP RS-68's from the launch pad on, since they cannot be air started, means a suboptimal ascent and cuts into payload capability compared to a pure first stage/second stage design optimized to use liquid fuels (kerosene first stage, hydrogen second stage). The RS-68's run out of fuel before the actual 'sweet spot' for EDS staging, but the core can't be made larger to hold enough fuel because it would be too heavy to lift off, and it would SEVERELY damage the performance capabilities... SO the EDS has to be larger (more dead weight through TLI) to carry enough propellant to complete the ascent to orbit, and still have enough to perform the TLI manuever. The extra dead weight from larger tankage on the EDS directly cuts into the TLI capability, reducing it pound for pound. The proposed EDS stage design using two seperate tanks and an intertank, instead of the common-bulkhead tanks used on the S-II and S-IVB stages, has a pathetic mass fraction (empty stage weight compared to propellant load) and is a poor design, basically a blunt instrument approach to performing TLI... it's not like common bulkhead tanks are some massive mystery that has yet to be solved... they've been around for nearly 45 years...

Von Braun and his team designed the Saturn V at a time when rocketry was basically in diapers and designed it all without benefit of computers, using slide rules and their considerable talent.

Here we are 50 years later with the benefit of all that experience and hindsight, with computer tools such as CAD and CFD that wasn't even dreamed of back then, and this kludge is the best we can do... which, unlike Saturn V, not only doesn't have any margins like Saturn V did, but cannot meet it's own performance requirements .

It's pretty pitiful IMHO... OL JR
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