Author Topic: Specific Impulse with Respect to SLS Booster type: Solid vs. Liquid.  (Read 31345 times)

Offline baldusi

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Have you taken in consideration the Height of the VAB? That's one big limitation. And I also think that with a commercial procurement approach the 10m tooling shouldn't be too expensive. But, this is a new paradigm for NASA, do exploration on a shoestring. Let's see what they learn. From what I've heard, they are on that "don't spend an extra cent than the strictly necessary for the 2017 IOC".
In my view, this is leading to some decisions that might come back as complications and extra costs on the long run. But I've also got the feeling that they think this is the last chance they have to build and design a LV. Thus, they have to be very aggressive on price and schedule, but very conservative on the innovation side. If they can't show built parts by 2015 and do testing by 2016, this program will most certainly be cancelled.

Offline Downix

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It's ironic that 10 m tooling costs are said to be prohibitive, since the USA's successful Apollo program used a rocket with two stages having the 10 m diameter. But, development costs, schedule constraints, and political factors are outside my field of expertise.

We have tools to make the existing ET based systems work, we do not for a 10m system.  A bird in the hand vs two in the bush.  Would 10m offer some advantages?  Sure.  However, would require an all new set of tools, which based on previous experience would take 3-4 years to get ready, so 3-4 years before the real work would begin.  On the other hand, using the existing ET tooling, the tools are already in place and ready to use, so can get to work right away.
chuck - Toilet paper has no real value? Try living with 5 other adults for 6 months in a can with no toilet paper. Man oh man. Toilet paper would be worth it's weight in gold!

Offline Rocket22

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Yes, the box that a LRB would have to fit to optimize performance during the first 120 seconds without breaking structural limits means that a higher ISP booster can burn longer than 120 sec while still having the same or very slightly better performance during the first 120 sec i.e. deltaV. A 311 ISP 3.4mlbf booster weighing 1.3mlb and burn time of ~145 sec would definitely system wise out perform a SRB. That first 120 sec is almost like it is set in stone. The more thrust an LRB booster has so that you can carry more propellant may increase the burn time more. There is a minimum thrust value for each ISP value that gives a SRB equivalent booster that burns 120 sec. By adding more thrust than that allows for carrying more propellant on the boosters allowing them to burn longer. Basically a 311 ISP booster would optimally be sized to burn at ~155 sec, thrust 3.6mlbf and 1.43mlb weight. A 280 ISP booster would be optimal for a burn time of ~138 sec, thrust 3.55mlbf and 1.39mlb weight (three 1.2mlbf Merlin 2 engines or 7 .5mlbf AJ-500’s). A 440 ISP booster would be optimal for a burn time of 218 sec, thrust 3.57mlbf and 1.4mlb weight. Because of the longer burn time they will deliver more overall performance i.e. payload than the SRB. Increasing the payload weight means an adjustment to the thrust or booster weight as well as burn time to optimize.

The SLS core design is a poor compromise for an efficient booster. We see that the CCB approach delivers better performance as well as lower overall costs. Using three boosters for the first stage burning for duration of 180sec, then a large LH2 US will greatly outperform SLS than using the same boosters on an SLS.

So if a LRB is designed and built, a tank stretch of the booster tanks, plus using 3 of them in line with a stretched SLS US would be about an equal performer to SLS. In fact a 5 engine 3 SLS core in line would work too, making the LRB’s a shortened core version of the SLS core itself or even the same length with crossfeed.

Geometrically, ten cylinders having diameters equal to that of the STS SRB's can be rung around the STS external tank (virtually a perfect fit I think). These diameters are sufficient for fairings that could contain out-rigged SSME's.  Thus, ten SSME's could provide the propulsion for the first stage of a three-stage HLLV that might be able to deliver 130 metric tons to low earth orbit. The diameter of such a first stage would equal the maximum diameter of the Shuttle. One caveat is that a full external tank could probably not supply enough LOX/LH2 for the first stage's optimum burn time (the SSME's would probably have to thrust at 109% of their design value). And, the cost of ten SSME's per launch might raise some eyebrows. But I wondered what others might think of this idea.

Offline 93143

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I'm told that the Michoud facility cannot handle a stage with a diameter greater than 10 m.

Last I heard, it couldn't even handle 10.  The overhead cranes weren't there back in the day; to do 33' now they'd have to either roll back to 1960s manufacturing techniques or raise the roof.

Offline spectre9

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If SSME engines are to be used for the SLS core, why not just eliminate the solid rockets, expand the diameter of the first stage to about 33', and fit it with 10 SSME engines?  Then, with 5 J2-type engines in a second stage and a single J2-type engine in a third stage, over 130 metric tons of payload could be delivered to low earth orbit with a tandem three-stage LOX/LH2 rocket weighing about 4.5 million lbs at liftoff. The large number of SSME engines needed in a heavy-lift program using this rocket should cause the unit cost of the engine to decrease significantly. From my limited perspective, I don't see how a rocket with two huge strap-on recoverable solid boosters could deliver more payload per dollar than an expendable LOX/LH2 three-stage rocket. I must be (and probably am) missing something.
From what I've read and been told, they had very specific time, budget and contractual constraints from the budget act. So they specifically went to the lowest development risk path, namely, ET tank core, RS-25D engines, Ares I solids, Ares I avionics, etc. The specifically wanted the least amount of new developments possible. Apparently, the 10m tooling was a huge cost. But more importantly, a huge development and certification risk.
I've seen a video, where they stated that a monolithic RP-1 first stage and H2/LOX second stage probably had the lowest operation cost, and best scalability. But the cost of development was big, but not ridiculous but they couldn't have anything flying before 2020.
From my humble position, the 2009/2011 ridiculous situation at NASA was caused by years of suspending investment in the fundamentals (big thrust RP-1 first stage engines and H2/LOX upper stage engines), and then doing too big and moving specs programs (Ares I/V). Can you believe that Griffin, in the recent Hearing stated, with a straight face, that the logical thing was his "presidential budget proposal" where he had 14B just for the Constellation program from 2013 onwards?
After that lesson, the ones that stayed at NASA learned the hard lesson and scoped programs for available budgets (around 3.5B of development money per year). Hence, the current SLS.

Is it a good idea to build a 10m core fully liquid fuelled rocket?

I'm not so sure about these SRBs getting the SLS to 130mt. Are they saying it can because that was it says in the authorisation act?

Are the costs of building new SRBs too high? I mean how much did the technology mature during the life of the shuttle?

Offline Rocket22

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If SSME engines are to be used for the SLS core, why not just eliminate the solid rockets, expand the diameter of the first stage to about 33', and fit it with 10 SSME engines?  Then, with 5 J2-type engines in a second stage and a single J2-type engine in a third stage, over 130 metric tons of payload could be delivered to low earth orbit with a tandem three-stage LOX/LH2 rocket weighing about 4.5 million lbs at liftoff. The large number of SSME engines needed in a heavy-lift program using this rocket should cause the unit cost of the engine to decrease significantly. From my limited perspective, I don't see how a rocket with two huge strap-on recoverable solid boosters could deliver more payload per dollar than an expendable LOX/LH2 three-stage rocket. I must be (and probably am) missing something.
From what I've read and been told, they had very specific time, budget and contractual constraints from the budget act. So they specifically went to the lowest development risk path, namely, ET tank core, RS-25D engines, Ares I solids, Ares I avionics, etc. The specifically wanted the least amount of new developments possible. Apparently, the 10m tooling was a huge cost. But more importantly, a huge development and certification risk.
I've seen a video, where they stated that a monolithic RP-1 first stage and H2/LOX second stage probably had the lowest operation cost, and best scalability. But the cost of development was big, but not ridiculous but they couldn't have anything flying before 2020.
From my humble position, the 2009/2011 ridiculous situation at NASA was caused by years of suspending investment in the fundamentals (big thrust RP-1 first stage engines and H2/LOX upper stage engines), and then doing too big and moving specs programs (Ares I/V). Can you believe that Griffin, in the recent Hearing stated, with a straight face, that the logical thing was his "presidential budget proposal" where he had 14B just for the Constellation program from 2013 onwards?
After that lesson, the ones that stayed at NASA learned the hard lesson and scoped programs for available budgets (around 3.5B of development money per year). Hence, the current SLS.

Is it a good idea to build a 10m core fully liquid fuelled rocket?

I'm not so sure about these SRBs getting the SLS to 130mt. Are they saying it can because that was it says in the authorisation act?

Are the costs of building new SRBs too high? I mean how much did the technology mature during the life of the shuttle?
As an engineer with experience in trajectory optimization and missile guidance theory but NOT in propulsion or structural engineering, it has always seemed to me that an all-liquid-propellant rocket could be made safer, less costly, and higher performing than a rocket employing solid rocket motors. The German rocket team probably realized this in the 1930's and 1940's as their developments advanced to the well-known A4 (later, V2) LOX/alcohol rocket.  Since we have extensive experience with liquid oxygen and liquid hydrogen as oxidizer/fuel propellants, and because this combination has such a relatively-high specific impulse, it seems logical to attempt to construct an all-LOX/LH2 rocket. Of course, the finest structural methods and materials might be required for the LH2 fuel tank, which will make for a long first stage which must sustain the expected in-flight propulsive and aerodynamic loads.

A problem with LOX/LH2 for the first stage of a rocket is getting the high thrust required to lift a heavy payload into orbit. Theoretically, ten SSME (RS-25) engines or eight RS-68 engines could be "ganged" (half of them out-rigged) to produce enough thrust for an 8.4-meter first stage, with propellant tanks having substantially more capacity than the Shuttle external tank does, to lift 130 metric tons to low earth orbit.  A second LOX/LH2 stage (8.4 m) with three J-2X engines and one RL-10 engine would then complete a rocket sufficient, theoretically, to lift a 130 metric ton payload to a 120 nautical mile orbit, with a liftoff weight much lower than that of the Saturn V.  Why the RL-10?  Well, it takes time to get to a 120 nautical mile altitude, and the J-2X engines can only be throttled about 20% (I think).  Now, the J2-X engines could be shut down and then re-ignited after a coast interval for the final orbit-injection burn.  But that final burn would be rather short (perhaps 12 seconds or so), and it would be better to have a lower-g burn in order to inject more accurately into orbit.  That's where the little RL-10 comes in.... to be used for the final injection burn… a long one (there are options other than an RL-10 for this, but the RL-10 is small and could probably be packaged with three J2-X engines within the 8.4 m diameter of the second stage).

Of course, there are countless candidate configurations for a LOX/LH2 heavy-lift launch vehicle.  A critical question is: Is it feasible to out-rig RS-25 or RS-68 engines onto an 8.4 meter core?


Offline Chris Bergin

Bumping some of the previous leading threads back to the top after the Space Policy HLV threads were moved into here.

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