Future of Prospective Raptor Upper Stage

[oldAtlas_Eguy]

Would the Raptor be bigger than the current second stage, as well as more efficient? Wondering if they could host a much larger second stage on Falcon Heavy. Apologies if already addressed.

Of course it could. The F9 US is optimized for F9, not FH.

A 5m diameter Raptor (or widebody F9 US) might make sense on an FH, *if* the performance is needed.

Could a 6 meter diameter or larger follow on upper stage be desirable to gain performance and/or increase payload volume.

It would look something like the proposed Atlas V Heavy with the long payload fairing

I went and checked the details on volumes etc for RP1/LOX vs LOX/LH2 and the factor is almost exactly 3 for the amount of volume needed by a LOX/LH2 stage over that of a RP1/LOX stage when holding the total propelant weight constant. Basiclly being able to deliver more delta V for the same ~stage weight due to higher ISP. You can create a stage that delivers ~the same delta V but it would have limited use beyond just being a more expensive replacement of the RP1 stage.

So a Raptor stage at 5m would be 1 to 1.5 times the tank length of the RP1 US.

So what is the most you could increase the delta V with a 5m stage? And could they significantly increase the LEO capability beyond 53mt?

Also, is there any news regarding development of Raptor? Is it only on paper, or are they prototyping?

Without changing the total stack weight of the FH by decreasing the propelant on the Raptor so for a given payload total system delta V is constant, the payload weight went up to ~66,000kg. If you add more propelant to the Raptor US increasing the total stack weight the 1st stage delta V decreases needing the US to produce more delta V just to get to the same total delta V. There are several tradeoff points for the propelant weights for the Raptor US under different destinations (total delta V required): LEO, GTO, Earth departure. I don't have enough data to model the FH 1st stage to do a first order tradeoff. Just estimates. The max LEO seems to be around 69,500kg. Where the Raptor stage has the same weight as the Rp1 stage.

P.S-total stack weight went up 20,000kg just to gain an additional 3,500kg of payload.

[oldAtlas_Eguy]

OOps I was pushing the tradeoff and found an error in the model.

For the same weight Raptor and RP1: Payload 70,300kg and increase in stack weight 17,000kg.

The extra 300kg is residual propelant not used.

I continued pushing the total stack weight up. When I got to a stage that was 3 times the length and weighed twice as much as the RP1 stage I stopped. The payload weight was approaching 80,000kg with an increase in stack weight of 50,000kg. Weight was increasing more rapid than payload was increasing. But if you needed it you could build a configuration that could put 80MT into LEO with Raptor. The liftoff acceleration will be lower but not yet significant change. The problem is structual loads on the first stage.An increase of 50,000kg is a 50% increase in loads requiring a redesign of the first stage. A addition of 17,000kg is only a 17% increase in loads one that would possibly not require a 1st stage redesign.

[DaveH62]

OOps I was pushing the tradeoff and found an error in the model.

For the same weight Raptor and RP1: Payload 70,300kg and increase in stack weight 17,000kg.

The extra 300kg is residual propelant not used.

I continued pushing the total stack weight up. When I got to a stage that was 3 times the length and weighed twice as much as the RP1 stage I stopped. The payload weight was approaching 80,000kg with an increase in stack weight of 50,000kg. Weight was increasing more rapid than payload was increasing. But if you needed it you could build a configuration that could put 80MT into LEO with Raptor. The liftoff acceleration will be lower but not yet significant change. The problem is structual loads on the first stage.An increase of 50,000kg is a 50% increase in loads requiring a redesign of the first stage. A addition of 17,000kg is only a 17% increase in loads one that would possibly not require a 1st stage redesign.

Would there be a limit on refiring the Raptor stage?

[starsilk]

OOps I was pushing the tradeoff and found an error in the model.

For the same weight Raptor and RP1: Payload 70,300kg and increase in stack weight 17,000kg.

The extra 300kg is residual propelant not used.

I continued pushing the total stack weight up. When I got to a stage that was 3 times the length and weighed twice as much as the RP1 stage I stopped. The payload weight was approaching 80,000kg with an increase in stack weight of 50,000kg. Weight was increasing more rapid than payload was increasing. But if you needed it you could build a configuration that could put 80MT into LEO with Raptor. The liftoff acceleration will be lower but not yet significant change. The problem is structual loads on the first stage.An increase of 50,000kg is a 50% increase in loads requiring a redesign of the first stage. A addition of 17,000kg is only a 17% increase in loads one that would possibly not require a 1st stage redesign.

Would there be a limit on refiring the Raptor stage?

the various 'teasers' that have been seen about Raptor describe it as a complex, staged combustion engine. usually the more complex engines require much more complex startup systems (it follows), so I'd expect it to be designed for the minimum number of restarts to complete the mission (ie 2 or 3).

[baldusi]

I think the minimum number would be five, to compete in capabilities with the Delta IV US. At least three would be the minimum to do an optimized GTO. What would be the minimum for GSO (plus US deorbit)?

[kevin-rf]

What would be the minimum for GSO (plus US deorbit)?

For disposal, Earth escape requires a lower Delta V. (*assuming no reuse)
Also, if the Raptor has the Delta V kit option, the disposal burn does not need to come from the raptor.

[spacetraveler]

Also, is there any news regarding development of Raptor? Is it only on paper, or are they prototyping?

Musk said at the AIAA event that they would be announcing a new engine development program soon. Does anyone know if this could be referring to Raptor or would it be some other engine?

There wasn't a lot of detail, other than it would be staged combustion.

[Jason1701]

Also, is there any news regarding development of Raptor? Is it only on paper, or are they prototyping?

Musk said at the AIAA event that they would be announcing a new engine development program soon. Does anyone know if this could be referring to Raptor or would it be some other engine?

There wasn't a lot of detail, other than it would be staged combustion.

Unknown to the public whether it is hydrogen or methane. Recall that most of what we think we know about Raptor comes from the infamous presentation.

[hkultala]

From another thread..

http://forum.nasaspaceflight.com/index.php?topic=28006.1395

SpaceX has bought some LH2 tanker cars from nasa, to be used in McGregor.

So this means raptor is not completely on hold and they are preparing for tests of raptor?



hmm, reasons what they could use raptor for..

1) To make a future version of Falcon Heavy equal SLS block 1 on capasity. But why?

2) To allow much heavier loads to be lifted towards mars with Falcon heavy. Mars is Elon's ultimate goal anyway.

3) To replace falcon heavy with falcon 9 on some bleo missions.




or..

what's the street price of rl-10 engine? what if they just buy some rl-10's to be used as third stage on falcon heavy (as an eds stage for mars missions) ?

[Robotbeat]

Buying some RL-10s is probably the least likely scenario, since even ULA is trying to move away from them because of their cost.

[FinalFrontier]

I doubt they would be buying RL-10s, it would not make sense given their history in up till now on engine development.

[arnezami]

SpaceX has bought some LH2 tanker cars from nasa, to be used in McGregor.

Incorrect. There is only 1 LH2 tanker car.

I wonder how much they are like the ones they already got... I see no reason why it should be used for LH2.

[Tcommon]

Buying some RL-10s is probably the least likely scenario, since even ULA is trying to move away from them because of their cost.

The RD-0146 is said to be an alternative, but I think P&W has marketing rights to it.

Unknown to the public whether it is hydrogen or methane. Recall that most of what we think we know about Raptor comes from the infamous presentation.

Hyundai has a methane engine they'd love to sell.

[hkultala]

Hyundai has a methane engine they'd love to sell.

ISP (sea) – 277 sec.
ISP (vacuum) – 321 sec.

So with it's current nozzle has WORSE vacuum isp than merlin 1c vacuum.

And even if they put a bigger nozzle, it's vacuum isp might only be slightly bigger than merlin vac's isp.

So I don't see spacex buying this when they have merlin 1dvac coming.

rl-10 or rd-0146 would have considerable better isp than either of these.

[sanman]

Even with Merlin1D getting so much more Isp out of the kerolox fuel than before, how can this overcome the fact LH2 is way the heck lighter than kerosene? It would seem that LH2's lightness puts it way far ahead, and that no kerolox engine can come close to it. (Unless someone finds a way to electrically accelerate kerosene) 

[peter-b]

The main problems with LH2 are that it's bulky (very low density compared to kerosene), which leads to enormous inflation in tank dimensions, and it is difficult to store and handle. For first stages, kerosene is pretty comparable with LH2 when the trade-off between achievable performance and practicality is considered.

[Kaputnik]

Even with Merlin1D getting so much more Isp out of the kerolox fuel than before, how can this overcome the fact LH2 is way the heck lighter than kerosene? It would seem that LH2's lightness puts it way far ahead, and that no kerolox engine can come close to it. (Unless someone finds a way to electrically accelerate kerosene) 

You've got to consider the whole system, not just the engine.
Specific impulse, propellant density, tankage mass, engine T:W, all play a part.

There is no 'one best' answer. It all depends on what you need to use the rocket stage for.

[Ben the Space Brit]

You've got to consider the whole system, not just the engine.

This is one of the reasons why I suspect that the next upper stage engine upgrade will be methane rather than hydrogen fuelled.  Hydrogen wold mean a lot of work and possibly a wide-body upper stage (5.2m, like the fairing).  You could just convert the RP1 tank in the existing F9US to store cryogenic methane, plug in the new engine and you're more-or-less ready to go.

[baldusi]

Even with Merlin1D getting so much more Isp out of the kerolox fuel than before, how can this overcome the fact LH2 is way the heck lighter than kerosene? It would seem that LH2's lightness puts it way far ahead, and that no kerolox engine can come close to it. (Unless someone finds a way to electrically accelerate kerosene) 

Liquid H2 is 14 times less dense than RP-1. But, since your O/F ratio is usually close to 6.0, and RP-1 uses something around 2.35, your effective propellant volume is "just" 3.0 times RP-1. You also need a lot more insulation, since H2 is liquid at around 40 Kelvin. So you not only need a three times bigger tank, but you need a lot more insulation.
Since the amount of LOX is not that different (H2 uses 20% more LOX volume per tonne of propellant), you have 6.25 times more H2 volume than RP-1 volume. In other words, you are replacing a barely insulated RP-1 tank with a six times bigger and heavily insulated H2. Also, remember that all your H2 piping has to allow six times more flow and be also heavily insulated than the equivalent RP-1 (roughly 2.5 times the diameter).
Which takes us to the engine. You have cryogenic H2 all through the engine, and thus have to keep huge temperatures differential. Luckily, H2 has a lot less pressure losses when used as the coolant. Regrettably, it also has 1/14 the mass, so you need to flow more volume of it (again, wider passages). Now you need to flow six times more fuel, your turbopump has to put six times more volume (but the lower density means that's not proportionally more power) and all the control valves need to be proportionally bigger. Remember, that getting the same pressure on a bigger diameter require proportionally more mass (mass growth cubically, diameter square). Thus, it's only natural that the H2 engines have worse T/W than an RP-1 engine.
On the other hand the specific impulse you can get with H2 is so ridiculously more than RP-1 than unless you are gravity losses dominated, you are better of with it.
The overall situation is that, in general, H2 is not better than RP-1 (or even CH4) for a first stage. But it's usually very advantageous for an Upper Stage. There's one caveat, though, for this. If you have a strategy of huge first stage and tiny upper stage (look into the EELV's proportion), it all depends a lot on how energetic you want you payload's orbit. Delta IV M is not that good at LEO, but shines on higher energy orbits.
Yet, you have to approach the problem from a whole system architecture. SpaceX's tries to get away with only a two stage system (or 2.5 in the Heavy version) and, basically, a single engine. From that point of view, a SC CH4 was the optimum to balance LEO and HEO. But then they are also betting on reusability. Also, they are sort of "volume limited". This might change everything.
Getting your upper stage from LEO is a lot easier than getting it from GTO in an usable form. What's more, a low T/W but high isp propulsion is incompatible with atmospheric flight. I'm speculating (heavily), that they will develop a tug. Or more than one. SEP propulsion gives excellent performance. So, what if they made an H2 tug to take satellites from LEO to GSO, and said tug had a SEP module to get back to LEO? What if they made an H2 tug to cross the Van Allen belt fast (humans or sensitive satellites) and a SEP tug to bring back the H2 tug or take said payloads to EML1/2 or even the Moon? If you are going to have reusability, you might find that anything but LEO is better done with tugs that never get back in the atmosphere. I'm speculating, obviously. But even the STS was planned for only LEO, so I believe that I'm not too far off.

[sanman]

Musk's main market is in LEO, with an expanding future market in GTO, and finally Musk's own ambitions are for Mars. I seriously doubt Musk would have gotten into the spacelaunch business if Mars was as barren as the Moon. LEO launch business is a stepping stone for him, and if that's all there was, then he wouldn't have gone into this line of work.

So it looks to me like he wants to get people over to his true love Mars ASAP. Raptor would be useful for that, because of its utility beyond LEO.

He said initial voyages to Mars would be 6 months, but that ultimately he'd like to cut it down to a month. I don't see how he's going to retire on Mars without a Raptor type upper stage to make Mars travel more energetic.