Possible SpaceX Vehicle Configurations based on recent (10/23) Raptor information

[meekGee]

So per information made available on 10/23, we now think that
a) Raptor is a Methane-burning 300-ton-force engine
b) Raptor will be tested at Stennis in "Early 2014"

The obvious questions now are:

- What vehicle will be built around it.
- What is the projected timeline.

(Previous threads were less constrained since there were several guesses as to the size of Raptor)

Relevant information on Raptor:
http://www.cochran.senate.gov/public/index.cfm/news-releases?ID=da0d793d-b2ec-4407-8b21-361505db2815
http://www.seattlepi.com/news/article/SpaceX-to-test-methane-rocket-engine-in-Miss-4919628.php
http://www.sunherald.com/2013/10/23/5053390/stennis-lands-spacex-another-rocket.html

[Joel]

I couldn't find the MCT speculation thread, but it was going kind of wild anyway.

We have a data point now - it seems like a 300 ton Methane engine is in the works.

What will it push?

300,000 kg is one big engine. I wonder how it will be transported.

[QuantumG]

hehe.. I sure hope that was a deliberate play on the stupid units because used to describe thrust here.

[Kabloona]

What will it push?

Only itself, if it weighs 300 tons. 

[meekGee]

hehe.. I sure hope that was a deliberate play on the stupid units because used to describe thrust here.

Since it's just an estimate, 2000 lbf vs. 2200 lbf. is no big deal..., and -f (kgf vs. kg) is a legit shorthand....

I am metric-native, and SI-natural....   But a "ton of thrust" is a good convenient measure.  It means it can lift a ton at T/W=1.

[Lars_J]

I couldn't find the MCT speculation thread, but it was going kind of wild anyway.

We have a data point now - it seems like a 300 ton Methane engine is in the works.

What will it push?

300,000 kg is one big engine. I wonder how it will be transported.

It would be smaller than the F-1 engine, so it could easily be transported by truck. Or do you mean how a large LV would be transported? Most likely built near the launch site.

[Lars_J]

Here is some fun with photoshop  - Imagining how a 7m LV powered by 7 Raptors might look next to F9 and FH. PURE speculation. 

[llanitedave]

A Raptor-powered single-core heavy that lifts payloads comparable to that of the FH would make re-use easier, since they wouldn't have to worry about the central core heading too far downrange.

[Zed_Noir]

Here is some fun with photoshop  - Imagining how a 7m LV powered by 7 Raptors might look next to F9 and FH. PURE speculation. 

Think a shorter 10 meter core with 11 Raptors plus 8 landing gear similar in size to the F9R units is also possible. Only the SpaceX CTO knows for sure. 

[GalacticIntruder]

One thing I was thinking of is SLS. SLS is an 8.4m Core. That means relevant tooling is already available to the industry. I think it would make sense for the SpaceX BFR to also adopt an 8.4m Core. Though that might not get them the Mars performance they want, so they could adopt a 10.1m Core, like the Saturn V. But my money is on 8.4m.

I also think single core is preferable to tri-core, given launch pad and complexity, but they could use smaller boosters for more performance.

[QuantumG]

SpaceX not making their cores in-house?

Wow, that'd be a change. If it turns out to be true, the BFR will have cost them more than any of us expected.

[Lars_J]

One thing I was thinking of is SLS. SLS is a 8.4m Core. That means relevant tooling is already available to the industry. I think it would make sense for the SpaceX BFR to also adopt an 8.4m Core.

Why? The only reason would be if SpaceX would subcontract the first stage to Boeing. Do you see that happening?

There is nothing magical or affordable about 8.4m tooling.

[beancounter]

One thing I was thinking of is SLS. SLS is a 8.4m Core. That means relevant tooling is already available to the industry. I think it would make sense for the SpaceX BFR to also adopt an 8.4m Core.

Why? The only reason would be if SpaceX would subcontract the first stage to Boeing. Do you see that happening?

There is nothing magical or affordable about 8.4m tooling.

The Chinese might disagree.

[Lars_J]

Why? The only reason would be if SpaceX would subcontract the first stage to Boeing. Do you see that happening?

There is nothing magical or affordable about 8.4m tooling.

The Chinese might disagree.

Care to clarify what you mean?

[beancounter]

Why? The only reason would be if SpaceX would subcontract the first stage to Boeing. Do you see that happening?

There is nothing magical or affordable about 8.4m tooling.

The Chinese might disagree.

Care to clarify what you mean?

Oh, there was an article out recently where they were delaying the large new launch vehicle because of quality issues surrounding the welding or construction of the vehicle.  It's much larger than their existing Long March hence my comment regarding 8.4m tooling - more to the 'magical' part than cost.  That's all.

[malu5531]

The obvious questions now are:

- What vehicle will be built around it.
- What is the projected timeline.

I did some calculations/guesstimates a while ago using the thrust now confirmed (650 klbf).

Rocket;
Diameter: 7.32m (twice F9)
Stack length (incl fairing): 94m
Stage 1 engines: 9 raptor (9x650 klbf, 332/360 Isp)
Stage 2 engines: 2 raptor vac (2x726 klbf, 390 Vac Isp)
Fairing: 10.4m diameter, 23m inner length (~1/3 FH cargo density)

Mass to LEO (200x200 @ 26 deg incl): 113mT
Note; Performance incl first stage boostback/reuse using >4000m/s residual delta-v post MECO

[Jason1701]

The obvious questions now are:

- What vehicle will be built around it.
- What is the projected timeline.

I did some calculations/guesstimates a while ago using the thrust now confirmed (650 klbf).

Rocket;
Diameter: 7.32m (twice F9)
Stack length (incl fairing): 94m
Stage 1 engines: 9 raptor (9x650 klbf, 332/360 Isp)
Stage 2 engines: 2 raptor vac (2x726 klbf, 390 Vac Isp)
Fairing: 10.4m diameter, 23m inner length (~1/3 FH cargo density)

Mass to LEO (200x200 @ 26 deg incl): 113mT
Note; Performance incl first stage boostback/reuse using >4000m/s residual delta-v post MECO

What would that be to TMI, and how about with three cores?

[Lars_J]

I did some calculations/guesstimates a while ago using the thrust now confirmed (650 klbf).

Rocket;
Diameter: 7.32m (twice F9)
Stack length (incl fairing): 94m
Stage 1 engines: 9 raptor (9x650 klbf, 332/360 Isp)
Stage 2 engines: 2 raptor vac (2x726 klbf, 390 Vac Isp)
Fairing: 10.4m diameter, 23m inner length (~1/3 FH cargo density)

Mass to LEO (200x200 @ 26 deg incl): 113mT
Note; Performance incl first stage boostback/reuse using >4000m/s residual delta-v post MECO

Nice speculation, a few notes:

1. Note that the 300t figure was vacuum thrust. Not sea level, which you probably should base your calculation on instead. I'm also not sure how you arrived at the 'confirmed thrust'. A rough unit conversion with something subtracted is not 'confirmed'.

2. Your spreadsheet also references a "75 mT Falcon 9 Heavy Raptor"... ? 

3. Overall your numbers do seem a tad bit optimistic. 113mT to LEO *with* reuse? I haven't done calculations, but that does sound off.

[gin455res]

In the slightly shorter term, how would a Falcon heavy with a 2 raptor core, oxgen cross-fed, expendible central core, with reusable kero-lox side boosters, carrying a superdraco kick stage perform for bleo missions?

[malu5531]

Nice speculation, a few notes:

1. Note that the 300t figure was vacuum thrust. Not sea level, which you probably should base your calculation on instead. I'm also not sure how you arrived at the 'confirmed thrust'. A rough unit conversion with something subtracted is not 'confirmed'.

2. Your spreadsheet also references a "75 mT Falcon 9 Heavy Raptor"... ? 

3. Overall your numbers do seem a tad bit optimistic. 113mT to LEO *with* reuse? I haven't done calculations, but that does sound off.

1) True
2) 75 mT FH is with a larger (more mass & wider) US with methane/raptor.
3) SpaceX fan; engine performance is on the upper limit, dry/wet ratios are low (4% first, 3.5% upper) - in line with musks 30:1 quote on booster ratio, adding a bit more for legs and interatage, and wider diameter should keep ratio low. Residuals are also low, such as 0.1% on US (only gasseous merhane and oxygen?)

I appreciate feedback on the calculations themselves; given my assumptions I believe this performance can be achievable (admittedly close to upper limit of possibility)