note: all figures here are in metric or SI units
*Extra note: all of this is speculation. Considering the early stages of the BFR project, only a fool would take this post as a cold hard truth. It’s speculation of what the SpaceX vehicle might realistically look like.*
So some time ago I had written a little blogpost about what a SpaceX Heavy Lifter using Raptor engines might look like, based on the latest info we had from SpaceX. However, recently, they updated the thrust figures for Raptor, and of course, higher thrust=bigger rocket=bigger payload. So I figured I’d update the figures, seeing as they got linked around quite a lot (the highest traffic source for this month was an /r/spacex thread).
In the post, I assumed the following:
– Stage 1:
– GLOW: 2452 tons
– Total propellant: 2305 tons
– Empty mass: 147 tons
– Thrust in vacuum: 40500 kN
– Thrust at lift-off: 35811 kN (3654 tons)
– Specific impulse: 363 vac, 321 sl, ~349 avg.
– Stage 2:
– GLOW: 582 tons
– Total propellant: 547 tons
– Empty mass: 35 tons
– Thrust: 4711 kN
– Isp: 380
However, the updated Raptor thrust figures are about 6915 kN for the first stage version and 8240 kN for the second stage. One thing that struck me here, however, is that the 705 tonne/6915 kN value is probably for sea level thrust, not for vacuum thrust. If the engine has the same mass flow rate for both vac and sea level, it would work out to the first stage version having a vacuum Isp of about 320 seconds, which is close to the previous value for the sea level Isp but not to the vacuum Isp. So if we assume this thrust is for sea level, the vacuum thrust becomes about 7820 kN.
If we scale up the rocket for these new thrust figures, and we assume 9 engines on the first stage we get the following:
**First stage**
- 4250 tons total mass
- Propellant mass: 3995 tons
- Empty mass: 255 tons
- Thrust at lift-off: 62235 kN (6345 tons)
- **Stage 2:**
- Full stage mass: 1018 tons
- Propellant mass: 957 tons
- Empty mass: 61 tons
- Thrust: 8240 kN (840 tons)
The Isps are the same as the last time. Again, I assumed a 10 ton fairing (similar to the 10 meter SLS fairing) that was separated around the same time that the first stage burned out. Using these figures, I got a useful LEO payload of anywhere between 260 and 293 metric tons, with 265 having a total ∆V of about 9500, and 295 assuming a ∆V of 9200 m/s. The methodology described here ended up giving me a payload of about 266 tons, giving the impression that the lower end of the spectrum is likely more realistic.
If this vehicle is sending payloads to Mars directly, the payload will be between 40 and 50 metric tons. If a third stage optimised for Mars is used, the payload becomes 80 tons to Mars if 266 tons to LEO is assumed and up to 90 tons if the payload of 293 tons is assumed. Hydrogen would increase this further, of course, but hydrogen is evil as we all know.
If the vehicle is made reusable, this payload drops pretty quickly of course. First stage reusability drops payload by about 30%, second stage reduces it by another 30%, or roughly halves the original payload (0.7*0.7=0.49). But even then, the vehicle would be able to deliver at least 130 metric tons into Low Earth Orbit.
I didn’t go into much detail on a multi-core variant, but some really quick estimates put the LEO payload at over 700 tons to LEO if made non-reusable… Fully reusable though, just like Falcon Heavy, the payload would go down very fast. As Falcon Heavy’s GTO payload drops to just 7 tons if all three first stages are reused, you shouldn’t expect this BFR to deliver much more than ~300-350 tons to LEO. Still, that’s huge beast.
To put the vehicle in perspective, the single-core vehicle has about 4952 metric tons of methane propellant. SLS Block 1B would carry about 1100 tons of liquid hydrogen and oxygen, and about 1300 tons of solid propellant, and Saturn V carried about 2744 tons of propellant in total. This machine would, if it really got built in this way, absolutely dwarf anything created before it.
Armchair Rocket Science 