Falcony Heavy with Raptor centre core and Merlin side boosters

[andyr]

Just wondered if this has been discussed anywhere ?

The advantages (forget the costs!) would be:

- Side boosters being Merlin can return to land.
- Central core is expendable
- F9 upper stage may move to Raptor anyway, so making the 1st stage Raptor also might not be too big a leap
- Central core has fewer Raptors (maybe 5?)
- Gains some Raptor flight time ahead of ITS

[macpacheco]

There's little point in considering a Falcon heavy type vehicle if putting raptor on the booster is a possibility.
With a 9 raptor booster + single raptor upper stage, assuming the stages are sized to efficiently use the higher thrust, we get a rocket that's more capable than FH with just 10 engines and no side boosters.
Each raptor has 2x the thrust plus there's the extra ISP.
Its notional performance to LEO would firmly invade SLS territory.
But far more important, it would have performance to launch 20 tons to GTO with full reuse, performance to launch a fully loaded red dragon to Mars or to the moon with at least booster reuse.
The key requirement to achieve such goals is to give up road transportability. It would have to be either constructed locally and shipped by barge. Something like a 5.2m diameter body to handle the lower density deep cryo metane.
With full reuse, road transportability becomes a much lower requirement.

If the goal is just to obtain more experience with Raptor in the easiest way, they just use it on the 2nd stage, where its been determined it will produce substantially higher performance even with the volumetric limitations of road transportability due to the higher ISP ruling in the 2nd stage.

In fact, I fully expect as soon as Raptor testing is concluded, SpaceX will either make a mini ITS (with multiple raptors in the upper stage, with a mix of vacuum and SL nozzles to enable retro landing), or a F9 on steroids with 6-9 raptors on the booster + a single raptor upper stage.

[andyr]

Hi - I was thinking the centre core would be of very similar size to the S1 current cores, with a reduced number of Raptors.

[Jim]

Hi - I was thinking the centre core would be of very similar size to the S1 current cores, with a reduced number of Raptors.

Not viable.  It would have less performance than the standard FH.  And mixing methane and Rp-1 doesn't make sense logistically.

[hkultala]

Just wondered if this has been discussed anywhere ?

The advantages (forget the costs!) would be:

- Side boosters being Merlin can return to land.

What is the point if you ado not care about cost?

- Central core is expendable

I don't see how this is "advantage".

- F9 upper stage may move to Raptor anyway, so making the 1st stage Raptor also might not be too big a leap

This is not advantage. You are just trying to claim  a "smaller disadvantage."
And there are NO released plans for raptor upper stage. It is simply not needed.

- Central core has fewer Raptors (maybe 5?)

5 raptors is still more expensive than 9 merlins. I don't see how than can have any "advantage".

- Gains some Raptor flight time ahead of ITS

... without recovering those engines and not getting to check them after flights. Just run them on test stand instead.

[gospacex]

There's little point in considering a Falcon heavy type vehicle if putting raptor on the booster is a possibility.
With a 9 raptor booster + single raptor upper stage, assuming the stages are sized to efficiently use the higher thrust, we get a rocket that's more capable than FH with just 10 engines and no side boosters.

I think thread originator meant F9-sized tankage - the same diameter as F9. Which means it can be built on the same factory with the same tooling as F9 cores are built, so it does make some economic sense.

9 Raptor booster can't be made with F9-sized tankage - 9 Raptors won't fit under it. One Raptor can.

[spacenut]

Someone calculated the nozzle size for a vacuum Raptor would be almost 5m wide.  That means it could not be used with the diameter of the existing upper stage of 3.7m.  Also only about 2-3 sea level Raptors could be used on the existing F9 core.  Too big.

It is not in the plans, but a mini-ITS say with a 6-7m core and 9 Raptors and one Vacuum Raptor upper stage would be a nice rocket.  It could match or exceed Falcon heavy with a single stick rocket.  But like I said, this in not in the plans. 

SpaceX is going straight for the ITS and the Big booster rocket. 

[Jim]

the same diameter as F9. Which means it can be built on the same factory with the same tooling as F9 cores are built, so it does make some economic sense.

No, it doesn't because it will have less performance and there is no plus side

[envy887]

Hi - I was thinking the centre core would be of very similar size to the S1 current cores, with a reduced number of Raptors.

Not viable.  It would have less performance than the standard FH.  And mixing methane and Rp-1 doesn't make sense logistically.

I'd like to see some math on the "lower performance" bit. The prop mass fraction is 21% better with kerolox, but Raptor's insane pressure and FFSC more than make up for it.

IF they hit Raptor's nominal specs, a 6100 kN, 2-Raptor booster with the exact same size and dry mass as the F9 booster, with the standard F9 upper stage, would put 3% more payload to LEO, by my math. And that's not accounting for savings on plumbing and thrust structure mass from having fewer engines, which you always say is significant.

With FH, making the center core methane is even better since as a "sustainer" type of stage it needs less thrust and higher ISP. A single-Raptor center stage (with same size/dry mass as FH core) would boost payload to LEO by around 10%.

I don't think the payload gain is in any way useful enough to be worth bothering the the logistical headaches, but it is at least not a net loss.

[tyrred]

Proving that the NSF Shepard is a healthy skeptic?

[Moderas]

@envy887 How would they land a 2-raptor (or 1 for that matter) booster? Why would SpaceX spend money designing a new vehicle that can not be reused, because it can not land, for a 10% performance gain over their current rocket which already has excess performance?

[Stan-1967]

Just wondered if this has been discussed anywhere ?

The advantages (forget the costs!) would be:

- Side boosters being Merlin can return to land.
- Central core is expendable
- F9 upper stage may move to Raptor anyway, so making the 1st stage Raptor also might not be too big a leap
- Central core has fewer Raptors (maybe 5?)
- Gains some Raptor flight time ahead of ITS

It has been discussed elsewhere. 
http://forum.nasaspaceflight.com/index.php?topic=41361.0

It is an interesting thought, the performance hit isn't that great if used as the propulsion for a non-reusable center sustainer core, but the reasons that it will never happen are very well founded.   

[envy887]

@envy887 How would they land a 2-raptor (or 1 for that matter) booster? Why would SpaceX spend money designing a new vehicle that can not be reused, because it can not land, for a 10% performance gain over their current rocket which already has excess performance?

How? Probably with 4 ten tonne gaseous methalox thrusters.

Why? It would have to be a while combination of factors pushing them in that direction, none of which are really likely for reasons already mentioned here.

[macpacheco]

- Gains some Raptor flight time ahead of ITS

... without recovering those engines and not getting to check them after flights. Just run them on test stand instead.

Let me add that IF (and thats a mighty big IF) SpaceX wanted Raptor flight time, that would be much better achieved like this:
Build F9 Raptor 2nd stage.
Make a variant with a sea level nozzle.
Launch and land that stage 10x at Spaceport America or any other rocket test range that would accept SpaceX. Given the development contract with USAF, I'd say SpaceX would be welcomed with open arms anywhere that has enough spare time for SpaceX.
Perhaps launch from Boca Chica and land at Boca Chica LZ.

Rocket engineering isn't done in the air, flying.
Rocket engines aren't flown before they have been tested extensively on the stand and have performed flawlessly.
Some might reply, oh but test stands don't replicate all flight conditions, true, but the likely hood of bona fide engine problems in flight is quite small.
There could be problems in the rocket-engine interface or in the interaction of a boatload of engines producing heat together. But such problems are likely specific to the design of a specific stage type rather than engine design/fabrication problem.

In conclusion, its quite unlikely that SpaceX needs Raptor flight time before ITS. The key is testing the whole rocket stage in its final configuration.
Even then, that testing is done with a full/partial mission duration test fire somewhere.

I will agree that it will be quite likely that once SpaceX builds either a full raptor rocket or a raptor 2nd stage to F9/FH that if that rocket configuration is fully reusable, it will undergo some flight testing program, perhaps 3 test launches, cause they can afford to do it with a fully reusable rocket that can be relaunched dozens of times.

[gospacex]

@envy887 How would they land a 2-raptor (or 1 for that matter) booster? Why would SpaceX spend money designing a new vehicle that can not be reused, because it can not land, for a 10% performance gain over their current rocket which already has excess performance?

They wouldn't land it.
This theoretical LV is intended to serve the heaviest payloads, which otherwise can't be launched by FH at all, thus they go to some competitor instead of SpaceX.
If the center core is single-raptor, expending it may be acceptable.

[DanielW]

I am pretty sure that falcon heavy will be SpaceX's last foray into kerbal style rocketry. By their own admission multi-core was much harder than anticipated.

[andyr]

Thanks for the (considered) replies.

My post was only for a discussion. My thoughts were that to have the central core Methalox (both 1st and 2nd stage), using current tankage size, but reusing Merlin powered side boosters, there could be some gains. Both in terms of payload and data.

I was expecting that this would have better performance TBH.

[Prettz]

More performance to do what?

[Brovane]

Just wondered if this has been discussed anywhere ?

The advantages (forget the costs!) would be:

- Side boosters being Merlin can return to land.
- Central core is expendable
- F9 upper stage may move to Raptor anyway, so making the 1st stage Raptor also might not be too big a leap
- Central core has fewer Raptors (maybe 5?)
- Gains some Raptor flight time ahead of ITS

Some simulations of performance with an FH and a Raptor S2 was run in L2.
http://forum.nasaspaceflight.com/index.php?topic=37599.300

[Steve G]

The only upgraded performance EM would want out of a Falcon Heavy would be to increase the Mars payload for those regular missions he mentioned. This would be best addressed by upgrading the upper stage and keeping everything else the same.  This is assuming he'd need more payload and if all he's launching is a Red Dragon that wouldn't be necessary. I don't see any changes to the Falcons once V5 is done.

The only way we would see a smaller iteration of the ITS is if they discover the challenges of building a full size ITS are too formidable in terms of construction or cost and compromise on a smaller design.

[TomH]

I'd like to see some math on the "lower performance" bit. The prop mass fraction is 21% better with kerolox, but Raptor's insane pressure and FFSC more than make up for it.

Methane has higher ISP, but RP-1 has higher ISP Density. Given the same volume, there is simply much more energy in RP-1. Given the same mass, there is more energy in the CH₄.

Math can help you compare, but math is not going to change the laws of chemistry.

[Rei]

I'd like to see some math on the "lower performance" bit. The prop mass fraction is 21% better with kerolox, but Raptor's insane pressure and FFSC more than make up for it.

Methane has higher ISP, but RP-1 has higher ISP Density. Given the same volume, there is simply much more energy in RP-1. Given the same mass, there is more energy in the CH₄.

You know, that makes me think of the possibility of going rather in the opposite direction, swapping out LOX for nytrox (say, 50%/-80°C), and ditching the tank pressurization system     Density is better than regular LOX (although not quite as good as subcooled), but beyond ditching (historically problematic) pressurization system's mass, it uses a significantly higher O:F ratio, and since the oxidizer is denser than the fuel....

Plus it's a lot nicer stuff in general. In addition to easier pressurization, it's also easier ignition, not nearly as cryogenic, and closer to RP-1 in temperature. The increasing nitrous fraction also increases chemical compatibility; unlike LOX, N2O is compatible with all common metals as well as many common organic materials. And the higher temperature means less likelyhood of cracking if one migrates to composites.  It retains its nontoxic/environmentally safe nature (N2O mixed with oxygen is a medical anesthetic and is safe for breathing for long periods of time), is much higher performance than pure N2O (although not as good as LOX), and unlike pure N2O, since most of the gas phase is oxygen, is not prone to spontaneous decomposition. It also burns cooler than pure LOX.  You could probably up the engine design pressure (and thus ISP and thrust) for a given level of corrosion.  And the higher temperature makes it more storable in space.

Hmm, what pressure are Falcon tanks pressurized to at launch?

[envy887]

I'd like to see some math on the "lower performance" bit. The prop mass fraction is 21% better with kerolox, but Raptor's insane pressure and FFSC more than make up for it.

Methane has higher ISP, but RP-1 has higher ISP Density. Given the same volume, there is simply much more energy in RP-1. Given the same mass, there is more energy in the CH₄.

Math can help you compare, but math is not going to change the laws of chemistry.

RP-1 has more energy, but its conversion to thrust in Merlin is far less efficient than Raptor's conversion of LNG. Delta-v is linear in exhaust velocity but logarithmic in mass ratio, so payload does not scale linearly with impulse density.

In a Falcon rocket body, the same volume of LNG at Raptor exhaust velocity has more delta-v than RP-1 at Merlin exhaust velocity:

2943 m/s * ln(460 tonne / 24 tonne) = 8688 m/s

3434 m/s * ln(384 tonne / 24 tonne) = 9516 m/s

You know, that makes me think of the possibility of going rather in the opposite direction, swapping out LOX for nytrox (say, 50%/-80°C), and ditching the tank pressurization system     Density is better than regular LOX (although not quite as good as subcooled), but beyond ditching (historically problematic) pressurization system's mass...

How are you going to press the RP-1 tank if you ditch the helium pressurization system?

[Jim]

This theoretical LV is intended to serve the heaviest payloads, which otherwise can't be launched by FH at all, thus they go to some competitor instead of SpaceX.
If the center core is single-raptor, expending it may be acceptable.

Wrong on all accounts. 
This theoretical LV is proof that rockets are not Legos and an meaningless exercise.

There are no " heaviest payloads" to be served in the first place.
Nor is there a competitor to lose it to.
Also, even if there was a payload or a competitor, there is no business case to build this vehicle.  It would require developing a seldom used core, upper stage and it would incompatible with existing launch sites.

Most of all, this configuration goes against all of Spacex's mantras.

It doesn't make any sense.

[Jim]

You know, that makes me think of the possibility of going rather in the opposite direction,

Bad idea and not feasible
a.  Lose too much performance
b.  does't solve the tank pressurization issue.  RP-1 still needs to be pressurized
c.  Tank pressure is less than 50 psi
d.  Loading the He slower or GOX autogenous pressurization are  better ideas to solve the perceived problem

[Rei]

You know, that makes me think of the possibility of going rather in the opposite direction,

Bad idea and not feasible
a.  Lose too much performance
b.  does't solve the tank pressurization issue.  RP-1 still needs to be pressurized
c.  Tank pressure is less than 50 psi
d.  Loading the He slower or GOX autogenous pressurization are  better ideas to solve the perceived problem

Thanks for the pressure (and note that concerning pressurants we're only talking about the oxidizer side). But could you clarify what you mean by "performance"? I'm going to guess specific impulse, rather than thrust?

Let's run the rocket equation for, say, for simplicity's sake, F9 s2, no fairing.  Because I can't be bothered to do the full staging setup 

Kerolox / 50% Nytrox-kerosene
ISP: 348 / ~328
O:F: 2,38 / ~4,5
(Data source: http://www.b14643.de/Spacerockets_2/United_States_1/Falcon-9/Merlin/index.htm)

Kerolox breakdown, from some quick googling:
RP-1: 32300kg
LOX: 75200kg
Total: 107500
(data source http://spaceflight101.com/spacerockets/falcon-9-ft/ - doesn't perfectly match the reported O:F, but close enough)

With assumed densified LOX density of 1230kg/mł and RP-1 at 860kg/mł:
RP-1: 37,56mł
LOX: 61,13mł
Total: 98,69mł

For Nytrox-Kerosene and a 4,5:1 O:F ratio:
 * RP-1: 17,94mł
 * Nytrox: 80,75mł

Thus, for nytrox density ~1190kg/mł:
 * RP-1: 16288kg
 * Nytrox: 96093kg
 * Total: 112381kg

We have a dry mass of 4000kg and say a 5000kg payload (9000). I don't know how heavy the oxidizer pressurant system is; I'll guess 500kg (feel free to correct with the actual number if you have it).  Our masses are thus:
 Kerolox: 116500
 Nytrox: 120881

Thus with ISPs of 348 and 328, respectively, we get the following dVs:
 Kerolox: 8732
 Nytrox: 9054

Nytrox wins by 322 m/s.  And by simplicity, and by chemical compatibility, and propellant storability, and ease of ignition, higher thrust, and everything else listed before.

Of course, the above is rendered virtually meaningless by the low feed pressure Falcons are designed for 

For fun, let's look at MON-10 at room temperature, which is around 3,5 bar (50 psi).  O:F ~4.5, so same as above.
Density on the MON is ~1425kg/mł  (http://www.dtic.mil/dtic/tr/fulltext/u2/a036741.pdf)
 RP-1: 15170kg
 MON: 115069kg
 Total: 130239 kg
With dry mass and payload:

Thus with ISPs of 348 and, say, 316, respectively, we get the following dVs:
 Kerolox: 8732
 MON: 8647

Kerolox wins by a smidge (85m/s)

Other additional factors.
 * On the downside, lower ISPs affect first stages more than upper stages.
 * On the upside, higher density means more thrust, which means lower gravity losses.
 * Also on the upside, lower flame temperatures means more potential for higher pressures (and even more thrust, plus a bit more ISP) without erosion.
 * On the downside, in the specific case of MON, it's more corrosive than Nytrox. And toxic.

But this is diverging from the original issue, which was that TomH is correct on - for a rocket of a fixed volume, switching to a higher ISP / lower density propellant is generally not a way to get better performance.

[Jim]

Nytrox wins by 322 m/s.  And by simplicity, and by chemical compatibility, and propellant storability, and ease of ignition, higher thrust, and everything else listed before.

No, does not win.  Nytrox is not going to be able to use a stock Merlin.   And you are ignoring tank mods for mixture ratio changes. And you cannot add in 500kg for pressurant system and also ignore the much much heavier Nytrox tanks.

It is not simple and there is no real difference in the ease of ignition (there never has been an issue with LOX RP-1 ignition)The pressures associated with it make the launch pad more complex. 

More thrust?  nonsense, you have no data to support that claim from a stock Merlin. 

[envy887]

... for a rocket of a fixed volume, switching to a higher ISP / lower density propellant is generally not a way to get better performance.

This entirely depends on the characteristics of the propellant and engine, and the operating environment of the rocket. There isn't a general case and each case has to be specifically analyzed.

In this case, assuming that Raptor+subcooled LNG has less payload to orbit than Merlin+ subcooled RP-1 appears to be incorrect. I have plotted in the attached image the bulk density required to achieve the Falcon 9 booster's delta-v with payload, assuming constant volume, stage dry mass, thrust, and outer mold line. Methalox Raptor is above the required I_sp to get the same performance as Merlin, while SSME-like hydrolox, even subcooled, is not.

Edit: this is discussed in some more detail here: https://forum.nasaspaceflight.com/index.php?topic=42302.0

[Rei]

Nytrox wins by 322 m/s.  And by simplicity, and by chemical compatibility, and propellant storability, and ease of ignition, higher thrust, and everything else listed before.

No, does not win.  Nytrox is not going to be able to use a stock Merlin.   And you are ignoring tank mods for mixture ratio changes.

Did you not read my very next sentence where I said it's a moot point, because it's not something you can just swap out into the system as-is? 

And you cannot add in 500kg for pressurant system and also ignore the much much heavier Nytrox tanks.

Look, do we really need to get into every last aspect of rocket design here, one at a time?  I point out that the pressurant tanks would be removed, which you already know. You point out that the propellant tanks would be heavier, which I already know. I point out that a higher feed pressure simplifies turbopump design requirements, more with increasing pressure, to the point that if it's high enough there's no turbopump required at all - which you already know.  You then follow up by point out that yes but the mass gained by doing doesn't compare to the mass increase from the tanks handling the higher pressures, which I already know. I then point out that the importance of dry mass to simplicity varies greatly depending on the role that a particular stage plays, with boosters and storable stages generally preferring simplicity at a cost of mass, which you already know. You then snipe at me because my calculations were based on an upper stage.  I then fire back that I explicitly pointed out that I just did an upper stage calculation for simplicity's sake...

Can we just skip this eminently predictable conversation?  Thanks.

It is not simple

Quite to the contrary, self-pressurizing stages are a hallmark of simplicity in rocketry, which is why they're so often used in spacecraft, where reliability takes top priority.

and there is no real difference in the ease of ignition (there never has been an issue with LOX RP-1 ignition)

Which is why spacecraft tend to use "easy to ignite" RP-1 rather than hypergolics, right?

The Merlin burns LOX/RP-1 in liquid phase.  Nitrox burns gas phase.  Gas phase means easier and faster mixing.  Eliminating uneven combustion stemming from uneven mixing is one of the biggest difficulties in rocket engine design programs. Gas mixtures tend to ignite without chugging and burn steadily. Liquids make you earn it.

More even mixing also tends to result in higher ISP - something which I didn't account for in the above.

More thrust?  nonsense

I'm going to assume that you didn't actually mean what you wrote, because do not believe that you wouldn't be aware that  higher propellant densities are associated with higher thrust for a given flow rate, all else being equal. A turbopump sweeps out a fixed volume per revolution and is designed to operate at a given RPM. The denser the propellant, the more energy is in that volume (again, all else being equal).

you have no data to support that claim from a stock Merlin.

And I'm not sure why you're inserting "for a stock Merlin" into this conversation, as if a stock Merlin could burn nitrox - or methane for that matter.  We don't live in a world where you just load a rocket with an arbitrary propellant that it wasn't designed for; there's always reengineering, to various degrees.  However, when it comes to engineering tradeoffs, if you want thrust, you want A) density, B) low corrosiveness, and C) low flame temperature, as the latter two are general limitations on how high of a pressure you can run without excessive erosion.

[JasonAW3]

You know, that makes me think of the possibility of going rather in the opposite direction,

Bad idea and not feasible
a.  Lose too much performance
b.  does't solve the tank pressurization issue.  RP-1 still needs to be pressurized
c.  Tank pressure is less than 50 psi
d.  Loading the He slower or GOX autogenous pressurization are  better ideas to solve the perceived problem

Jim, with respect, you forgot one...

e.  The need to load four different liquids, (LOX, LHe, Methane and RP-1) all of which use different cryogenic requirements, over complicates the fuel loading procedures and exponentially increases the risk factor for a mishap.  Better to read that as a disaster.

[Jim]

1.  Quite to the contrary, self-pressurizing stages are a hallmark of simplicity in rocketry, which is why they're so often used in spacecraft, where reliability takes top priority.


2.  Which is why spacecraft tend to use "easy to ignite" RP-1 rather than hypergolics, right?

The Merlin burns LOX/RP-1 in liquid phase.  Nitrox burns gas phase.  Gas phase means easier and faster mixing.  Eliminating uneven combustion stemming from uneven mixing is one of the biggest difficulties in rocket engine design programs. Gas mixtures tend to ignite without chugging and burn steadily. Liquids make you earn it.


3.  I'm going to assume that you didn't actually mean what you wrote, because do not believe that you wouldn't be aware that  higher propellant densities are associated with higher thrust for a given flow rate, all else being equal. A turbopump sweeps out a fixed volume per revolution and is designed to operate at a given RPM. The denser the propellant, the more energy is in that volume (again, all else being equal).

1.  Huh?  so often used?  Name one launch vehicle stage that does use it(and doesn't include autogenous or weapons). Much less, name one spacecraft using "self-pressurizing" propellants.   

2.  We aren't talking about spacecraft or hypergolic.  We are talking about LOX/RP-1 vs Nitrox/RP-1.  What "easier" or "simpler" ignition system would be employed with Nitrox/RP-1 that couldn't be used with LOX/RP-1?

3  That is the issue, all else is not equal.  There is no guarantee that any engine can deliver the numbers you are stating.   And certainly not a stock Merlin. 

The OP was taking an "existing" engine with its propellants and constraining it to an existing vehicle length and diameter. 
There is no such Nitrox/RP-1 engine that exists (in any form) to support your claims that Nitrox/RP-1 core would be better performing than a Merlin powered LOX/RP-1 core.

[Jim]

We don't live in a world where you just load a rocket with an arbitrary propellant that it wasn't designed for; there's always reengineering, to various degrees.  However, when it comes to engineering tradeoffs,

Correct and the difference of  Kerolox: 8732 m/s vs  Nytrox  9054 m/s with your unsupported engine estimates* are too close to make any sort of judgment on.  Especially to say Nytrox "wins', when it doesn't even enter the race.

*Kerolox / 50% Nytrox-kerosene
ISP: 348 / ~328

328 is too high of an estimate compared to RP-1 theoretical vs actual.


 

[gospacex]

Nor is there a competitor to lose it to.

ULA thinking: "There is no competition (now), so why bother trying to do better?"

[Prettz]

Nor is there a competitor to lose it to.

ULA thinking: "There is no competition (now), so why bother trying to do better?"

This would be doing it worse, since it would explode costs.

If you're thinking of a "build it and they will come" type deal regarding payload capabilities, Falcon Heavy already is that thing. Why take something that's already a good idea and ruin it?

[Jim]

Nor is there a competitor to lose it to.

ULA thinking: "There is no competition (now), so why bother trying to do better?"

Another inane comment.  You are making yourself look like an ass

There is no competition because there is no payload or launch vehicle in this class

[envy887]

There is no competition because there is no payload or launch vehicle in this class

There's Orion, if SpaceX wants FH to compete with SLS. There's a host of other issues, obviously, but right now FH simply can't launch Orion BLEO due to lack of performance.

And Lunar Dragon mission, which has at least 2 paying customers - that was definitely a case of extra LV performance opening up a new market. More performance might open up the lunar orbital (vs. flyby) market.

[Jim]

on, if SpaceX wants FH to compete with SLS. There's a host of other issues, obviously, but right now FH simply can't launch Orion BLEO due to lack of performance.

And Lunar Dragon mission, which has at least 2 paying customers - that was definitely a case of extra LV performance opening up a new market. More performance might open up the lunar orbital (vs. flyby) market.

No, Orion is not a viable payload for FH.   And Lunar Dragon is not a case for this. 

The OP configuration doesn't do anything to help those anyways, which is the point of this thread, and not idiotic posts like
https://forum.nasaspaceflight.com/index.php?topic=42722.msg1666741#msg1666741

Anybody thinking that this configuration is good for Spacex doesn't know a thing about Spacex.  This configuration is a kludge and Spacex would not give it one second of consideration.

[macpacheco]

Rei there's one issue that shoots down your idea. How can you guarantee that RP1 and N2O will mix uniformly such that mixture ratios can be assured ?
That alone can kill your idea.
For instance, RP1 is a non polar molecule by simple virtue of being a quite large CH2-CH2-CH2-... chains. It would like to mix up with other hydrocarbons it doesn't like to mix with water.
Wouldn't N2O be at least a far more polarized molecule that would like to mix with water instead ?
Wouldn't this mean that even if a perfect mix of RP1 and N2O is initially loaded into the tank, wouldn't N2O and RP1 start to form RP1 rich areas and N2O rich areas ?
That's just one issue. I'm no rocket engineer.
I learned that in high school chemistry.
There are other similar issues beyond polar/non polar that I'm 100% unsure, but let be shoot them anyways, like solubility limits. It might make one of the two materials float to the top of each other.
Are you a chemist ? I'm certainly not.

[hkultala]

Let's run the rocket equation for, say, for simplicity's sake, F9 s2, no fairing.  Because I can't be bothered to do the full staging setup 

You are totally ignoring the fact that heavier second stage means first stage cannot lift it to so high velocity. The staging would happen something like 0.1 km/s earlier when the secnd stage is 5 tonnes heavier.

You are also taking the "328s" isp for nytrox from a hat, no source for that?

What if the true isp with the nytrox would be about 300s instead?

And no, for zillion other practical reasons this makes no sense at all.

[Lemurion]

SpaceX always optimizes for cost.

I fail to see how either the OP's mix of raptor and merlin cores or shifting to a Nitrox/RP-1 mixture is going to be cheaper for SpaceX to implement than the company's current plans.

It's like the Mini-ITS idea - I can see the benefits of having one but I can't see SpaceX building one because I don't see them building something that's not on their critical path. SpaceX doesn't need Mini-ITS so why spend the money to build one?

[gin455res]

SpaceX always optimizes for cost.

I fail to see how either the OP's mix of raptor and merlin cores or shifting to a Nitrox/RP-1 mixture is going to be cheaper for SpaceX to implement than the company's current plans.

It's like the Mini-ITS idea - I can see the benefits of having one but I can't see SpaceX building one because I don't see them building something that's not on their critical path. SpaceX doesn't need Mini-ITS so why spend the money to build one?

Has anyone Modeled a sustainer core with 2-4 over-expanded + 1 standard central merlins?


Could this reach orbit without an upper-stage?


Re-entry may be interesting as the angle of entry will be much flatter, the core will initially be weightless, and the grid fins will enable a high-lift entry, surfing a very narrow shockwave.  (Can a core become a true hypersonic glider with the greater control authority of the larger grid fins.)


How would the 'wing-loading' of a 3-5 engined core compare to a capsule, or the shuttle orbiter?


Current entries the booster weighs nearly all 100% its empty mass and it is coming in much more steeply.

[Lars-J]

Has anyone Modeled a sustainer core with 2-4 over-expanded + 1 standard central merlins?

How and where would those over-expanded Merlins fit?

[rakaydos]

Has anyone Modeled a sustainer core with 2-4 over-expanded + 1 standard central merlins?

How and where would those over-expanded Merlins fit?

Take out 2 regular merlins to fit 1 slightly overexpanded one between them. Probably have to kill the center gimbal too.

[watermod]

Despite describing the Block 5 as the final version of the Falcon 9, Shotwell hinted at the possibility of a future version of the rocket that could use the Raptor engines designed for transportation to Mars. Shotwell said Raptor, a liquid methane and oxygen engine for SpaceX’s interplanetary spaceship, has undergone “many dozens of tests” and is progressing well.

“The original idea for those engines were to serve as a propulsion system for the big Mars system, but we are looking at the utility of it on the Falcon program,” she said.

http://spacenews.com/spacexs-final-falcon-9-design-coming-this-year-two-falcon-heavy-launches-next-year/?utm_content=buffered598&utm_medium=social&utm_source=twitter.com&utm_campaign=buffer

[whitelancer64]

Despite describing the Block 5 as the final version of the Falcon 9, Shotwell hinted at the possibility of a future version of the rocket that could use the Raptor engines designed for transportation to Mars. Shotwell said Raptor, a liquid methane and oxygen engine for SpaceX’s interplanetary spaceship, has undergone “many dozens of tests” and is progressing well.

“The original idea for those engines were to serve as a propulsion system for the big Mars system, but we are looking at the utility of it on the Falcon program,” she said.

http://spacenews.com/spacexs-final-falcon-9-design-coming-this-year-two-falcon-heavy-launches-next-year/?utm_content=buffered598&utm_medium=social&utm_source=twitter.com&utm_campaign=buffer

To me, "looking into the utility" is a way of saying "we're confirming that it's not going to work."

Much like NASA "looked into the possibility" of putting crew on the EM-1 mission but concluded that it wasn't feasible.

[gin455res]

Has anyone Modeled a sustainer core with 2-4 over-expanded + 1 standard central merlins?

How and where would those over-expanded Merlins fit?

Take out 2 regular merlins to fit 1 slightly overexpanded one between them. Probably have to kill the center gimbal too.

Yes, or just remove engines 2, 4, 6, 8  and move 1,3,5,7 inboard slightly and add a larger bell
perhaps for the three engine version have two pumps feeding four bells, plus the central standard engine