Author Topic: Possible SpaceX Vehicle Configurations based on recent (10/23) Raptor information  (Read 225977 times)

Offline Robotbeat

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Any multicore configuration that is not arranged in a single line ( side by side) makes it impossible to do horizontal integration.

R7 is a five multicore configuration not inline, with horizontal integration.
Those are strap on boosters, not cores of equal size to the main core. The issue is how the whole stack will be supported without putting excessive strain on the cores or connections.
...actually, at the bottom of the rocket, the boosters /area/ the same size as the main core. Much of each of their dry mass (engines, plumbing, vernier rockets) is pretty close to the same scale as the center core, it's just the tank sizes which are different.


...but anyway, I agree with those who think that it's likely the next rocket that SpaceX develops after Falcon Heavy will probably be a large, TSTO rocket with no sideboosters. I don't think they'll need side-boosters, and I think that the scaling laws point to the idea of not making a 300mT monster with big side boosters. You'd destroy the pad with the thermal and acoustic environment of launch, and support buildings would have to be quite far away, reducing the economic effectiveness.
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Offline dglow

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I don't think they'll need side-boosters, and I think that the scaling laws point to the idea of not making a 300mT monster with big side boosters. You'd destroy the pad with the thermal and acoustic environment of launch, and support buildings would have to be quite far away, reducing the economic effectiveness.

That's interesting. The pad issues cited... are these the scaling laws you refer to? Or are there additional, say, flight profile considerations?

Offline Robotbeat

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I don't think they'll need side-boosters, and I think that the scaling laws point to the idea of not making a 300mT monster with big side boosters. You'd destroy the pad with the thermal and acoustic environment of launch, and support buildings would have to be quite far away, reducing the economic effectiveness.

That's interesting. The pad issues cited... are these the scaling laws you refer to? Or are there additional, say, flight profile considerations?
Nothing that couldn't be engineered around. I mean, you COULD still do it. It's just that acoustic energy becomes an important consideration as you scale your rocket bigger and bigger. These things give off lots of energy.

I mean think about it! If a 300mT launch vehicle has 3x the lift-off thrust of Saturn V (i.e. about 100MN) with an exhaust velocity of 3km/s or so, you're talking about over 300 tons of propellant expended before you even clear the tower. If you include the first 20 seconds, that's .7kiloton of propellant. Since kerosene makes up about 30% of that, but has about 10 times the energy density of TNT, that's equivalent to the amount of energy of a 2kt nuclear weapon! And that's energy that has to be dealt with by the pad somehow (diverted, shielded against, etc). That's a fourth of a Nagasaki right there.
« Last Edit: 01/18/2014 08:25 pm by Robotbeat »
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Offline AncientU

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I don't think they'll need side-boosters, and I think that the scaling laws point to the idea of not making a 300mT monster with big side boosters. You'd destroy the pad with the thermal and acoustic environment of launch, and support buildings would have to be quite far away, reducing the economic effectiveness.

That's interesting. The pad issues cited... are these the scaling laws you refer to? Or are there additional, say, flight profile considerations?
Nothing that couldn't be engineered around. I mean, you COULD still do it. It's just that acoustic energy becomes an important consideration as you scale your rocket bigger and bigger. These things give off lots of energy.

I mean think about it! If a 300mT launch vehicle has 3x the lift-off thrust of Saturn V (i.e. about 100MN) with an exhaust velocity of 3km/s or so, you're talking about over 300 tons of propellant expended before you even clear the tower. If you include the first 20 seconds, that's .7kiloton of propellant. Since kerosene makes up about 30% of that, but has about 10 times the energy density of TNT, that's equivalent to the amount of energy of a 2kt nuclear weapon! And that's energy that has to be dealt with by the pad somehow (diverted, shielded against, etc). That's a fourth of a Nagasaki right there.
A 27 engine 'monster' would produce close to 2x the lift-off thrust of SLS and 2.3x that of Saturn V.  No one expects catastrophic damage from SLS launches nor did we experience them from Saturn V.  Doubling the lift-off thrust will about double the thermal energy, maybe the acoustic energy, too.  But comparison to instantaneous thermal energy release and resulting blast wave of a nuclear weapon is nonsense.
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Offline Robotbeat

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Pad operations like what we saw for Shuttle or Saturn V operations are not exactly simple. There's a lot of pad damage that occurs.
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Offline JasonAW3

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Big issues with N1 is that they didn't properly ground-test the first stage, not necessarily the number of engines.

I'm not really sure that it would have really mattered much anyway.  The plumbing of the 36 engines was through one core and used a central plumbing system. Anything affecting that main line would affect the entire rocket.  One engine blowing up would tend to chain react through the whole system.

     In this case, we're talking 4 core stages, with 7 engines engines each, with pumps and plumbing for each engine.  From what I could see of the N-1's setup, it appeared that it used one main pump system for all 36 engines.  I'm most likely wrong on this point, but from the few diagrams I've seen, that's what it looks like was being done.

     As they say, 'keep it simple stupid!'

Jason

« Last Edit: 01/19/2014 12:35 am by JasonAW3 »
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Offline AncientU

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Pad operations like what we saw for Shuttle or Saturn V operations are not exactly simple. There's a lot of pad damage that occurs.
Not saying they are simple at all.  Just that 4 launches in seven months (Dec 68-July 69) indicates that the pad(s) were not ground zero (totally destroyed).  The release of energy is awesome, and no doubt damages much in the vicinity, but doubling it is not catastrophic or a reason to not build a 300mT launcher as previously implied.  I would suspect that if a F9 super-size is built with Raptors, a 3-core heavy version won't be far behind.

BTW, the fame trench at Pad39A is too narrow (18m) for such a launcher, so considerable rebuilding will be needed.  Will be interesting to see what they build to start with at Boca Chica.
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Offline Hyperion5

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I don't think they'll need side-boosters, and I think that the scaling laws point to the idea of not making a 300mT monster with big side boosters. You'd destroy the pad with the thermal and acoustic environment of launch, and support buildings would have to be quite far away, reducing the economic effectiveness.

That's interesting. The pad issues cited... are these the scaling laws you refer to? Or are there additional, say, flight profile considerations?
Nothing that couldn't be engineered around. I mean, you COULD still do it. It's just that acoustic energy becomes an important consideration as you scale your rocket bigger and bigger. These things give off lots of energy.

I mean think about it! If a 300mT launch vehicle has 3x the lift-off thrust of Saturn V (i.e. about 100MN) with an exhaust velocity of 3km/s or so, you're talking about over 300 tons of propellant expended before you even clear the tower. If you include the first 20 seconds, that's .7kiloton of propellant. Since kerosene makes up about 30% of that, but has about 10 times the energy density of TNT, that's equivalent to the amount of energy of a 2kt nuclear weapon! And that's energy that has to be dealt with by the pad somehow (diverted, shielded against, etc). That's a fourth of a Nagasaki right there.

Robotbeat, while I admit your words brought a smile to my face, let's at least keep the hand-wringing accurate.  :)  Here's what you should be wringing your hands about:


1.The table with the Falcon XX parameters is attached.
2.Falcon XX Heavy has the payload of 295 metric tons (watch the table). Thus I removed heat shields and reduced the mass of a head fairing.
3 . Expenses of speed on to become to Mars make 4000 m/s.
4 . When using aerodynamic braking in the atmosphere of Mars the mass of payload in a Martian orbit will make:
- 100 metric tons when using for dispersal from Earth of the hydrogen engine
- 70 metric tons when using the methane engine.






It's has ~96% more thrust than an SLS Bloc I (in thrust) and 115% more than a Saturn V, so AncientU is correct that the difference is not 3X.  It is however much more efficient, lifting 321% more to LEO than an SLS Bloc I and 150% more than the Saturn V.  While it would undoubtedly set the record for rocket decibels and tear up the pad, that it probably not the major problem for a prospective Raptor-powered HLV family.  Reusability would probably be the biggest challenge.  At the very least the heat shields would trim payload to at most ~288 mt on the tri-core and 97.7 mt on the single core. 

Does anyone have a reasonable estimate what landing legs on all three cores would cost in terms of payload?  Do we even know what the Falcon 9R's legs mass? 

Offline Robotbeat

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I don't know, but I really doubt we're going to be dealing with a tricore MCT. Also, why not ~7m diameter (as we've been hearing) instead of 9.8m? The upper stage looks tiny with such a wider diameter. Also, I'm not so sure they're looking at a full 100mT payload.

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Offline Hyperion5

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I don't know, but I really doubt we're going to be dealing with a tricore MCT. Also, why not ~7m diameter (as we've been hearing) instead of 9.8m?


That 9.8 m diameter was a result of Dmitry Vorontsov curbing the dry mass as much as possible.  He later said he could reduce the ullage in the oxygen tanks by about 17-18% by going to a 9.2 m diameter.  Whether or not that's the ideal (best compromise between dry mass & aerodynamic drag) is not entirely clear.  Dmitry mentioned it would take a lot more work to determine that ideal diameter.  If not a tri-core with 27 Raptors, then what?  How else do you plan to get a 50 mt MCT to Mars in only 1-2 launches?  Put all 27 engines on a single core? 

The upper stage looks tiny with such a wider diameter. Also, I'm not so sure they're looking at a full 100mT payload.

That large diameter makes launching wider MCTs easier, and that makes aerodynamic braking easier.  As for the 100 mt payload, I'm not sure I understand what you mean.  Do you mean to LEO or to Mars? 

Offline Robotbeat

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The upper stage /IS/ the MCT. The MCT is an upper stage, lander, and ascent vehicle. That fits what Musk tweeted. Also has the fewest pieces for a complete architecture. Also, if the MCT enters Mars' atmosphere similar to how the F9 upper is to renter Earth's, then it will probably enter sideways.
« Last Edit: 01/19/2014 07:03 am by Robotbeat »
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Offline DJPledger

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I don't know, but I really doubt we're going to be dealing with a tricore MCT. Also, why not ~7m diameter (as we've been hearing) instead of 9.8m?


That 9.8 m diameter was a result of Dmitry Vorontsov curbing the dry mass as much as possible.  He later said he could reduce the ullage in the oxygen tanks by about 17-18% by going to a 9.2 m diameter.  Whether or not that's the ideal (best compromise between dry mass & aerodynamic drag) is not entirely clear.  Dmitry mentioned it would take a lot more work to determine that ideal diameter.  If not a tri-core with 27 Raptors, then what?  How else do you plan to get a 50 mt MCT to Mars in only 1-2 launches?  Put all 27 engines on a single core? 

SpaceX will want to develop a larger version of the Raptor if they go for a single core to launch a 50mt MCT to Mars to keep the number of engines to 9 on the core. The 300tf at vac. Raptor currently under development is most likely to be used on a fully reusable single core LV to replace FH with 9 on it's 1st stage and 1 powering it's 2nd stage.

Offline MP99

I have a query about autogenous pressurisation. It feels like something that must have been discussed before, but limited search via Tapatalk...

I don't think this can work for an upper stage or MCT due to long loiter times required, so this is in relation to the first stage (SL Raptor) only.

I've seen a discussion that Russians might consider using a burner at the tank head to generate pressure gasses - basically H2O & CO2 at pretty high temperatures when your inputs are kerosene & lox. AFAICT, they would get away with this because there is little time for the propellants to be contaminated with water & CO2 ice during the short burn. Any crystals that do form would be small enough to go through the turbopumps if it came to that.

Could something similar be done with Raptor, but taking a tapoff from the pre-burners? One of the benefits of FFSC is that you have a high temperature stream of mostly-O2 that could pressurise the O2 tank, and mostly-CH4 that could pressurise the CH4 tank. This depends on the temperatures being feasible for this - I know they're very high, but ISTR not insanely high. Ducting that all the way to the tank heads would not be trivial. Perhaps the temperatures could be moderated by diverting some unburnt tank liquid into the return gas stream?

The O2 pressurant would have CO2 & H2O contaminants, while I'd guess the CH4 pressurant would add some CO (and OH?) to the mix?

Of course, one of the issues is what would happen to those gasses when they mix with cryogenic liquids in the tanks during coast phase before boostback & landing? Would ice crystals reduce the lifetime of the turbopumps for reuse?

I also wonder how the pressure would evolve during coast phases? Pressurants will cool in contact with the prop, but some prop will be boiled off in cooling / freezing them.

Cheers, Martin

Offline butters

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I have a query about autogenous pressurisation. It feels like something that must have been discussed before, but limited search via Tapatalk...

I don't think this can work for an upper stage or MCT due to long loiter times required, so this is in relation to the first stage (SL Raptor) only.

I've seen a discussion that Russians might consider using a burner at the tank head to generate pressure gasses - basically H2O & CO2 at pretty high temperatures when your inputs are kerosene & lox. AFAICT, they would get away with this because there is little time for the propellants to be contaminated with water & CO2 ice during the short burn. Any crystals that do form would be small enough to go through the turbopumps if it came to that.

Could something similar be done with Raptor, but taking a tapoff from the pre-burners? One of the benefits of FFSC is that you have a high temperature stream of mostly-O2 that could pressurise the O2 tank, and mostly-CH4 that could pressurise the CH4 tank. This depends on the temperatures being feasible for this - I know they're very high, but ISTR not insanely high. Ducting that all the way to the tank heads would not be trivial. Perhaps the temperatures could be moderated by diverting some unburnt tank liquid into the return gas stream?

The O2 pressurant would have CO2 & H2O contaminants, while I'd guess the CH4 pressurant would add some CO (and OH?) to the mix?

Of course, one of the issues is what would happen to those gasses when they mix with cryogenic liquids in the tanks during coast phase before boostback & landing? Would ice crystals reduce the lifetime of the turbopumps for reuse?

I also wonder how the pressure would evolve during coast phases? Pressurants will cool in contact with the prop, but some prop will be boiled off in cooling / freezing them.

Cheers, Martin

I think MCT will use an integrated vehicle fluids (IVF) approach using APUs (fuel cells or combustion engines) to generate power and exhaust gas for refrigeration and pressurization of the propellant tanks. It seems like their architecture will require them to be able to transport LOX and LCH4 over the long haul between Earth and Mars (either with the passengers or in a separate tanker), so active refrigeration would be required.

Offline baldusi

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ULA's IVF uses an internal combustion engine. The N-1 used a small turbine to generate electricity. So any engine is fair since it could provide electricity and pressurizant.

Offline go4mars

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Long before I had read the term autogenous pressurization, I mentioned the idea on one of these threads of using heat collected in the "regeneratively cooled nozzle" to phase change some of the liquid in each the methane and oxygen tanks (thereby increasing their pressure).

I talked about how the idea was based on the falcon 1 flight 3 thrust transient, and mentioned the desirable idea of low-rate gaseous combustion instead of liquid, but using the same hardware. 

This was in supposition that the tanks needed a low flow-rate "plume cushion" to reduce g's when re-entering the atmosphere and increases turgor pressure of the tanks.  (hot sauce viagra method for re-entry).

No one seemed to like the idea but me.  I still think it's a 2 birds with one stone solution but as always, would welcome some feedback.
« Last Edit: 12/24/2014 04:16 am by go4mars »
Elasmotherium; hurlyburly Doggerlandic Jentilak steeds insouciantly gallop in viridescent taiga, eluding deluginal Burckle's abyssal excavation.

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