Still want an SSTO? Your move.

[sevenperforce]

Prompted by a lengthy discussion over on @KerikBalm's "near-future scifi" thread, I created a rather beefy Excel table to model the performance and fuel fraction of a nuclear-thermal turbocharged ramrocket engine, to get an idea of what might be possible for SSTO applications. The spreadsheet worked so well that I expanded it out to provide fuel fractions for a wide range of engine types using various propellant combinations

If you want to build an SSTO, you've gotta be able to fit into these fuel fractions. No getting around it. These are, within error, the absolute minimum fuel fractions you'll need to make orbit. If the table says a given fuel fraction is 85.5%, then you've gotta fit your engines, tanks, airframe, and payload (plus margins and recovery hardware) into the remaining 14.5% of GLOW. Note that I didn't even include the values for a "true" airbreather (e.g., turbojet/ramjet/scramjet), because the thrust losses of combusting the airstream while still trying to accelerate make net performance far, far worse than even a pure rocket design.

With no further ado, here's the table!

Fuel fractions for SSTO

	Precooled	Turbocharged	Air-augmented	Rocket only
NTR (H2O)	N/A	72.3%	73.2%	83.1%
NTR (LH2)	47.9%	48.9%	49.7%	63.7%
Hydrolox	79.8%	80.7%	81.5%	89.0%
Methalox	85.5%	85.9%	86.6%	92.3%
Kerolox	N/A	89.1%	89.7%	94.3%

Visually:

Each SSTO ascent profile is constructed around a base assumption of 7.8 km/s to orbit, plus 750 m/s of gravity drag and 750 m/s of air drag.

The NTRs are Tantalum Halfnium Carbide pebble-bed reactors operating at slightly over 4400 K; basically the best thing we could actually build with modern tech.

Assumed specific impulses:

• NTR (H₂O): 469 s at SL, 555 s in vacuum
• NTR (LH₂): 820 s at SL, 971 s in vacuum
• Hydrolox: 366 s at SL, 452 s in vacuum
• Methalox: 334 s at SL, 382 s in vacuum
• Kerolox: 282 s at SL, 348 s in vacuum 

Additional notes...

Rocket only. This is provided mostly for comparison. It is assumed that a rapid climb is used to leave the atmosphere as soon as possible, with peak specific impulse and zero drag being reached around 2 km/s. This design invariably has the worst fuel fraction but allows the highest TWR engines. As mentioned above, altitude compensation is assumed; specific impulse climbs linearly to 2 km/s before plateauing at the vacuum value.

Air-augmented. This is an optimally designed intake shroud/duct with area behind the duct for expansion. A 22% static thrust boost is assumed due to pressure entrainment. Thrust augmentation reaches 100% around Mach 1.75, then begins to drop around Mach 6, decreasing to zero at the exhaust velocity of the actual engine. Aerodynamic drag is higher, at 800 m/s, and is spread out over a larger range of airspeeds, with vacuum specific impulse being reached much later. This represents a large fuel fraction gain for a fairly modest decrease in engine TWR. Specific impulse starts at slightly higher than the vacuum isp, then rises rapidly before dropping gradually.

Turbocharged. This adds a single-stage compressor fan to the shroud inlet. The increase in static thrust is substantial, allowing a physically smaller engine, but there is only a fractional improvement in net fuel fraction, as the fan is only useful to about Mach 2. Because the fan intake is more demanding than a ram intake, the design and vehicle integration may prove to cost significantly more in dry mass than it saves. However, the turbocharger can (in principle) be used alone as a ducted fan for a hover-light landing, which is a nice and very efficient advantage if you're looking for that (e.g., with an integrated-cabin manned launch vehicle). Aerodynamic drag increases to 850 m/s. 

Precooled. This is the design employed by SABRE, albeit without combustion of the airstream. Precooling the intake air allows the compressor fan to operate up to Mach 5 with hydrogen and Mach 3 with densified methane, with a corresponding improvement in fuel fraction. Water and kerosene are excluded, for obvious reasons. Dry mass penalty will be hefty, though.

For those who REALLY want to nerd out, here are the specific impulse curves, using hydrolox as an example:

Spoiler

And here are samples of the drag curves with respect to velocity:

Spoiler

 

Edited April 14, 2017 by sevenperforce

[sevenperforce]

As an example of how this might be used...

...here is a table showing the dry mass of a putative "Dragon 3" methalox SSTO crew shuttle using an air-augmented rocket only and a crew of 7. Dry mass is provided based on what tankage ratio and static engine TWR you can achieve. Vehicle dry mass includes the mass of the crew cabin and airframe, given in kg.

So if you can achieve a tankage ratio of 21 and an air-augmented static TWR of 53.7, you can have an SSTO crew shuttle with a dry vehicle weight of less than 15 tonnes, roughly the dry weight of a two-seat supersonic jet fighter.

[wumpus]

The link to the "air breather" makes a few assumptions that are designed to kill any thoughts of a true air breather.  From these slides it looks like the team behind the X-43 (the airbreather with easily the highest top speed) planned to build a mach ~7 first stage (with expendable second stage) as their first orbit-capable launcher.  That seems to be about the limit of their scramjet to accelerate things (it could maintain mach ~9.8 but with only the smallest acceleration).  Note that the falcon9 (in recovery mode) seems to leave the atmosphere at roughly the same speed, so even pure rockets still have drag issues at such speeds.

The whole idea of an SSTO has even more issues when competing with rockets that recover "spent" stages.  While you don't have to re-integrate the stages back together, you will have more maintenance issues as the entire rocket has to go back through orbital re-entry velocity, instead of the ~10% the recoverable multi-stage rocket experienced the brunt of reentry.  Not only does the rocket equation work against you, but the economics of launching are hardly the slam dunk that brought the SSTO dream into existence.  Don't forget that total fuel bill will always be higher, making the SSTO a possible brief window of low costs in going into space, followed by being a fuel guzzler that might become unlaunchable if fuel costs become significant (certainly large percentages of the population won't be able to get into space until  either the fuel costs dominate, or some other magic means to space exists (see metastable metalic mercury and/or space elevators).

[sevenperforce]

52 minutes ago, wumpus said:

The link to the "air breather" makes a few assumptions that are designed to kill any thoughts of a true air breather.  From these slides it looks like the team behind the X-43 (the airbreather with easily the highest top speed) planned to build a mach ~7 first stage (with expendable second stage) as their first orbit-capable launcher.  That seems to be about the limit of their scramjet to accelerate things (it could maintain mach ~9.8 but with only the smallest acceleration).  Note that the falcon9 (in recovery mode) seems to leave the atmosphere at roughly the same speed, so even pure rockets still have drag issues at such speeds.

Trouble with the X-43 is that the scramjet can only be used as a second stage, at best. In the current version, it is the second stage from an air-drop, meaning you would effectively need four stages to reach orbit.

Now, if you could come up with a way to do a parallel air-drop launch with a liquid rocket boosting the whole vehicle to scramjet ignition speeds, accelerating to Mach 7, and then separating so the rocket carries the payload to orbit and the scramjet is recovered, that might work.

Anyway, this project was as much intended as an SSTO-excluding product as an SSTO-enabling one. If someone fields an SSTO design with lower fuel fractions than the table calls for, you can know immediately that it's not going to work.

It also shows, within error, the relative savings of going partial-airbreathing/RBCC over pure rocket.

[Steel]

55 minutes ago, wumpus said:

...see metastable metalic mercury...

I'm assuming you mean metastable metallic hydrogen

[shynung]

There's a concept for a pulsed NTR reactor here that can boost effective reactor power output by, obviously, pulsing the reactor. From Atomic Rockets section on it:

Quote

The pulsed nuclear thermal rocket is a type of solid-core nuclear thermal rocket concept developed at the Polytechnic University of Catalonia, Spain and presented at the 2016 AIAA/SAE/ASEE Propulsion Conference. It isn't a torchship but it is heading in that direction. Thanks to Isaac Kuo for bringing this to my attention.

As previously mentioned, solid core nuclear thermal rockets have to stay under the temperature at which the nuclear reactor core melts. Having your engine go all China Syndrome on you and shooting out what's left of the exhaust nozzle in a deadly radioactive spray of molten reactor core elements is generally considered to be a Bad Thing. But Dr Francisco Arias found a clever way to get around this by pulsing the engine like a TRIGA reactor. The engine can be used bimodally, that is, mode 1 is as a standard solid-core NTR (Dr. Arias calls this "stationary mode"), and mode 2 is pulsed mode.

Pulse mode can be used two ways:

---

Direct Thrust Amplification: Garden variety solid core NTRs can increase their thrust by shifting gears. You turn up the propellant mass flow. But since the reactor's energy has to be divided up to service more propellant per second, each kilogram of propellant gets less energy, so the exhaust velocity and specific impulse goes down.

But if you shift to plus mode along with increased propellant mass flow, the reactor's effective energy output increases. So you can arrange matters in such a way that each kilogram of propellant still gets the same share of energy. Bottom line: the thrust increases but the specific impulse is not degraded.

---

Specific Impulse Amplification: This is really clever. For this trick you keep the propellant mass flow the same as it was.

In a fission nuclear reactor 95% of the reactor energy comes from fission-fragments, and only 5% come from prompt neutrons. In a conventional solid-core NTR the propellant is not exposed to enough neutrons to get any measurable energy from them. All the energy comes from fission fragments.

But in pulse mode, that 5% energy from neutrons could be higher than the 95% fission-fragment energy in stationary mode. The difference is that fission fragment energy heats the reactor and reactor heat gives energy to the propellant. And if the reactor heats too much it melts. But neutron energy does not heat the reactor, it passes through and directly heats the propellant.

The end result is that in pulse mode, you can actually make the propellant hotter than the reactor. Which means a much higher specific impulse than a conventional solid-core NTR which running hot enough to be right on the edge of melting.

Thermodynamics will not allow heat energy to pass from something colder to something hotter, so it cannot make the propellant hotter than the reactor. But in this case we are heating the propellant with neutron kinetic energy, which has zippity-do-dah to do with thermal transfer.

The drawback of course is that the 95% fission-fragment energy is increased as well as the neutron energy. The important point is by using pulsing you can use an auxiliary cooling system to cool the reactor off before the blasted thing melts, unlike a conventional NTR.

Apparently Dr. Arias' paper claims the pulsed NTR can have a higher specific impulse than a fission fragment engine. I am no rocket scientist but I find that difficult to believe. Fission fragment can have a specific impulse on the order of 1,000,000 seconds.

---

How Does It Work?

TRIGA reactor have what is called a large, prompt negative fuel temperature coefficient of reactivity. Translation: as the nuclear fuel elements heat up they stop working. It automatically turns itself off if it gets too hot. Technical term is "quenching."

Which means you can overload it in pulses. The TRIGA is designed for a steady power level of 100 watts but you can pulse the blasted thing up to 22,000 freaking megawatts. It automatically shuts off after one-twentieth of a second, quickly enough so the coolant system can handle the waste heat pulse.

 

[Rakaydos]

15 hours ago, sevenperforce said:

Trouble with the X-43 is that the scramjet can only be used as a second stage, at best. In the current version, it is the second stage from an air-drop, meaning you would effectively need four stages to reach orbit.

Now, if you could come up with a way to do a parallel air-drop launch with a liquid rocket boosting the whole vehicle to scramjet ignition speeds, accelerating to Mach 7, and then separating so the rocket carries the payload to orbit and the scramjet is recovered, that might work.

Anyway, this project was as much intended as an SSTO-excluding product as an SSTO-enabling one. If someone fields an SSTO design with lower fuel fractions than the table calls for, you can know immediately that it's not going to work.

It also shows, within error, the relative savings of going partial-airbreathing/RBCC over pure rocket.

What about rail-launched scramjets? We can build "cars" that break the sound barrier- how hard would it be to build a track designed to bring a large scramjet (and it's non-airbreathing upperstange) to operational speed at ground level?

[Zeiss Ikon]

3 hours ago, Rakaydos said:

What about rail-launched scramjets? We can build "cars" that break the sound barrier- how hard would it be to build a track designed to bring a large scramjet (and it's non-airbreathing upperstange) to operational speed at ground level?

Going to pull a Heinlein and run it up the side of Pike's Peak?  Mauna Kea might be a better choice (higher peak, lower base, and closer to the equator) but there's that pesky native religious preserve near the peak...  The issue is, the scramjet won't work well at sea level air density; it needs to be up high (lower stratosphere, by preference) so thrust can overcome drag.  So, you run your linear accelerator launcher up the side of a mountain, let the vehicle coast up until it's near level at at 10-15 km altitude, and then start up your scramjet.

Seems to me there are a bunch of pretty impressive mountains virtually on the equator in Colombia and Ecuador.  Colombia has a coast on the (north)west, you could start the rail at convenient sea level and near a port (Cartagena is likely to be a good choice).  Oh, good luck cleaning up the criminal element there.  Maybe the frequent, extremely loud (because large machine going very-very fast) sonic booms will cause them to plant their coca trees somewhere else...

[sevenperforce]

23 hours ago, shynung said:

There's a concept for a pulsed NTR reactor here that can boost effective reactor power output by, obviously, pulsing the reactor. From Atomic Rockets section on it:

 

Yeah, a pulsed NTR is basically as close to a torchship as we can realistically get with modern tech. Neutron flux would make it rough for manned missions. Though with water as remass, the neutron absorption rate would be pretty good....

4 hours ago, Zeiss Ikon said:

Going to pull a Heinlein and run it up the side of Pike's Peak?  Mauna Kea might be a better choice (higher peak, lower base, and closer to the equator) but there's that pesky native religious preserve near the peak...  The issue is, the scramjet won't work well at sea level air density; it needs to be up high (lower stratosphere, by preference) so thrust can overcome drag.  So, you run your linear accelerator launcher up the side of a mountain, let the vehicle coast up until it's near level at at 10-15 km altitude, and then start up your scramjet.

A scramjet simply can't get meaningful acceleration when it is needed most. Intake drag is too high.

[Zeiss Ikon]

1 minute ago, sevenperforce said:

A scramjet simply can't get meaningful acceleration when it is needed most. Intake drag is too high.

Hence the linear accelerator launcher.  Of course, if you're going to build one of those, a pretty small upgrade in G limits would let you launch direct to LEO, at least, needing only a circularizing motor (a few hundred m/s?) along with the payload.  Payload fraction (as launched) 80%?  Run the rail from Cartagena to a suitable peak in the northern Andes, and you could launch at less than 5 G even direct to orbit (possibly as low as 3 G).  You could probably run the thing on hydro and wind power, without even the need for a fission power plant.

Think of it as Musk's Hyperloop with a set of partial pressure doors at one end.  Doors, say a couple dozen a couple miles apart, open sequentially as the vehicle approaches, but never enough open at once to significantly repressurize the tube, and you'd use a conventional two-gate airlock to load the vehicle at the bottom end.

[hypervelocity]

this is great, I have been trying to model a SABRE spaceplane in RSS+RO with the b9 engines and only reached ≈4.5km/s before fuel depletion - I will evaluate if my design fits the aforementioned table & tweak it if it does not.

it would be great to tie this thread to users' actual attempts at building these in RSS/RO !  

Edited April 16, 2017 by hypervelocity

[wumpus]

21 hours ago, sevenperforce said:

A scramjet simply can't get meaningful acceleration when it is needed most. Intake drag is too high.

Thanks to the rocket equation you "need it most" the faster you are going.  But adding an additional stage to something that isn't going to break mach 7 is pretty silly for even near-modern tech: you only would consider doing such if fuel costs were coming into play.  Maybe after a decade or so of Musk's "hundred launches per booster" hyperbole, fuel costs may look significant enough to consider this.

[sevenperforce]

7 minutes ago, wumpus said:

21 hours ago, sevenperforce said:

A scramjet simply can't get meaningful acceleration when it is needed most. Intake drag is too high.

Thanks to the rocket equation you "need it most" the faster you are going.  But adding an additional stage to something that isn't going to break mach 7 is pretty silly for even near-modern tech: you only would consider doing such if fuel costs were coming into play.  Maybe after a decade or so of Musk's "hundred launches per booster" hyperbole, fuel costs may look significant enough to consider this.

Indeed. Mach 7 is only 2.4 km/s...within the range of typical staging velocity for a TSTO. Why would you build two stages (booster + scramjet) to do something that a single rocket-powered stage can do?

Not to mention that a scramjet needs to run on LH₂ in order to provide any meaningful performance, so fuel costs are still going to be more expensive than something like kerolox or methalox.

21 hours ago, Zeiss Ikon said:

21 hours ago, sevenperforce said:

A scramjet simply can't get meaningful acceleration when it is needed most. Intake drag is too high.

Hence the linear accelerator launcher.

No, you misunderstand my point. In order to get to orbit rapidly, a launch vehicle needs to have more and more acceleration as it gains velocity. A rocket goes from 1 km/s to 2 km/s very, very quickly. But for a scramjet (or any airbreather), it accelerates less and less as it gains velocity, because the momentum of the intake air is subtracted from its total thrust. So even though the fuel consumption for the total thrust is still very good, the effective specific impulse (impulse to the vehicle itself) plummets.

This is often called the "airbreather's burden" and it places a hard limit on the upper speed of an airbreather; it also means that the final portion of acceleration barely crawls along.

19 hours ago, hypervelocity said:

this is great, I have been trying to model a SABRE spaceplane in RSS+RO with the b9 engines and only reached ≈4.5km/s before fuel depletion - I will evaluate if my design fits the aforementioned table & tweak it if it does not.

Note that a SABRE is a "true" airbreather, meaning its effective specific impulse is far higher than a (comparatively simple) air turboramrocket. I was modeling for a ramrocket only rather than a crossfed system like SABRE. It has to deal with correspondingly higher intake drag, though. I can model it easily enough if I have a way of calculating net thrust with respect to velocity.

When it comes to Skylon itself, of course, bets are off. I used the assumption of a TWR>1 with a rocket-like ascent profile. Since Skylon uses lift to counteract gravity drag, I would need to model it differently, using lift-to-drag ratios relative to airspeed. Straightforward enough, but a different beast.

[shynung]

On 16/4/2017 at 11:03 PM, sevenperforce said:

Yeah, a pulsed NTR is basically as close to a torchship as we can realistically get with modern tech. Neutron flux would make it rough for manned missions. Though with water as remass, the neutron absorption rate would be pretty good

Well, that's the whole point. A pulsed reactor would try to heat propellant by neutron flux rather than through the reactor structure.

Though, you can always tow the payload behind the rocket.

Also solves the problem of getting stuff/people on and off the rocket, if using a flexible connector.

Also worth considering that if propellant cost is less of a concern than reactor fuel, it's useful to have a nuclear rocket motor than has a variable Isp. This allows a lower-output reactor to have a useful TWR at the cost of propellant fraction, which can ultimately save on nuclear fuel costs.

[sevenperforce]

8 minutes ago, shynung said:

Well, that's the whole point. A pulsed reactor would try to heat propellant by neutron flux rather than through the reactor structure.

Though, you can always tow the payload behind the rocket.

Yes, though that doesn't work so well for launching from Terra Firma.

One really cool idea is the NTER, or nuclear-thermal-electric rocket. The problem with NTRs is that if you heat them too much, they melt. You can get around this in a number of ways (like pulsed reactors), but NTER is a particularly ingenious method.

Instead of depending on thermodynamic transfer, the NTER uses a closed-cycle helium coolant loop to drive a pair of counter-rotating turbines. These turbines produce a rotating magnetic field which induces extremely high resistance heating due to the interacting electrical and magnetic fields. This converts the excess heat energy from the reactor into mechanical energy, which is then converted back into heat via electromagnetic interaction, increasing propellant temperature without overheating the reactor.

[shynung]

10 minutes ago, sevenperforce said:

Yes, though that doesn't work so well for launching from Terra Firma.

Launching, I think. would be fine. It's like a plane towing a glider. Atmo entry, on the other hand, might by tricky; I'm thinking about reeling in the payload section at entry, then reeling it back out during descent and landing. Then, after payload section touches down, the engine section must maneuver to touch down a safe distance away.

15 minutes ago, sevenperforce said:

One really cool idea is the NTER, or nuclear-thermal-electric rocket. The problem with NTRs is that if you heat them too much, they melt. You can get around this in a number of ways (like pulsed reactors), but NTER is a particularly ingenious method.

Instead of depending on thermodynamic transfer, the NTER uses a closed-cycle helium coolant loop to drive a pair of counter-rotating turbines. These turbines produce a rotating magnetic field which induces extremely high resistance heating due to the interacting electrical and magnetic fields. This converts the excess heat energy from the reactor into mechanical energy, which is then converted back into heat via electromagnetic interaction, increasing propellant temperature without overheating the reactor.

This is a pretty clever way to end up with an exhaust stream hotter than the reactor. It'd be somewhere between pure NTR and pure nuclear-electric rocket, in terms of ISP, I think.

[sevenperforce]

8 minutes ago, shynung said:

Launching, I think. would be fine. It's like a plane towing a glider. Atmo entry, on the other hand, might by tricky; I'm thinking about reeling in the payload section at entry, then reeling it back out during descent and landing. Then, after payload section touches down, the engine section must maneuver to touch down a safe distance away.

I...wouldn't trust it. Aerodynamics seem like they would be ridiculously screwy.

9 minutes ago, shynung said:

This is a pretty clever way to end up with an exhaust stream hotter than the reactor. It'd be somewhere between pure NTR and pure nuclear-electric rocket, in terms of ISP, I think.

They were projecting 900 seconds at 100 kN of thrust with a TWR of 0.3; useful for a Mars transfer but not so much for launch.

[shynung]

23 minutes ago, sevenperforce said:

I...wouldn't trust it. Aerodynamics seem like they would be ridiculously screwy.

Yeah, it's basically a whole other game. Not impossible, but pretty tricky. Even the HELIOS program whose picture I posted proposed launching it reeled-in, using a chemical booster.

23 minutes ago, sevenperforce said:

They were projecting 900 seconds at 100 kN of thrust with a TWR of 0.3; useful for a Mars transfer but not so much for launch.

That kind of surprised me. 900 seconds is the typical NTR ISP. I was expecting higher ISP at lower TWR.

Edited April 17, 2017 by shynung

[sevenperforce]

12 minutes ago, shynung said:

Yeah, it's basically a whole other game. Not impossible, but pretty tricky. Even the HELIOS program whose picture I posted proposed launching it reeled-in, using a chemical booster.

That kind of surprised me. 900 seconds is the typical NTR ISP. I was expecting higher ISP at lower TWR.

Their goal was not so much to push ISP to the limit as it was to allow a lower operating temperature for the nuclear core, to mitigate problems found with NERVA.

[shynung]

Since we're talking about nuclear Earth-based SSTOs, I agree with @KerikBalm at their thread that the nuclear lightbulb is the way to go. Barring a pulsed-solid core rocket, it's one of the few reactors that can run hot enough for a sufficiently torch-like performance on almost all situations.

[sevenperforce]

14 minutes ago, shynung said:

Since we're talking about nuclear Earth-based SSTOs, I agree with @KerikBalm at their thread that the nuclear lightbulb is the way to go. Barring a pulsed-solid core rocket, it's one of the few reactors that can run hot enough for a sufficiently torch-like performance on almost all situations.

To be honest, I'm a little bit in love with the TaHfC liquid-uranium pebble-bed design with water as propellant. 550 seconds of vacuum specific impulse and a TWR of 260:1 is really, really impressive. Supposedly the 260:1 includes the mass of the reactor, the turbopump, the nozzle, and the gimbal. That's a 12,000 kN thruster. Supposedly efficiency can go up a little more if you make the thruster smaller.

 

[Nothalogh]

5 hours ago, shynung said:

A pulsed reactor would try to heat propellant by neutron flux rather than through the reactor structure.

I hadn't thought of that approach, that is ingenious.

[Exoscientist]

 In regards to the kerosene propellant fraction notice SpaceX has already reached that with their Falcon 9 series. So the only thing needed to get the SSTO would be to swap out the Merlin for a higher ISP engine.

 Such engines do exist, created by the Russians for example.

  Bob Clark

[sevenperforce]

6 hours ago, Exoscientist said:

 In regards to the kerosene propellant fraction notice SpaceX has already reached that with their Falcon 9 series. So the only thing needed to get the SSTO would be to swap out the Merlin for a higher ISP engine.

 Such engines do exist, created by the Russians for example.

Unfortunately the higher-ISP kerolox engines are ORSC, which makes them much heavier for their thrust than the Merlin engines.

The RD-180 boasts a SL specific impulse of 311 seconds, and kerolox has a maximum theoretical vacuum isp of 353 seconds. Using these numbers decreases fuel fraction to 93.7%, which is not a huge improvement but it's something. Note that this doesn't account for increased gravity drag resulting from slower fuel consumption.

So we have the Merlin 1D with a TWR of 183:1 and a fuel fraction of 94.3%, and the RD-180 with a TWR of 78.44 and a fuel fraction of 93.7%. Fuel fraction calculations were made with the assumption of a LV TWR of 1.3:1, so with the Merlin 1D this requires 0.77% of GLOW to be engine mass, leaving 5.53% of GLOW for payload, tankage+structure, and margins. For the less powerful but more efficient RD-180, we have 1.66% of GLOW as engine mass, leaving 4.64% of GLOW for payload, tankage+structure, and margins. So switching from the Merlin 1D to the RD-180 is a worse tradeoff, even before you factor in the increased gravity drag of the lower fuel consumption. It's possible that the RD-180 could have improved TWR/isp tradeoff using a variable mixture ratio as used in the F1 engines on the Saturn V first stage, but it would still have some difficulty.

On the pad, the Falcon 9 first stage carries 4.23 tonnes of engine, 411 tonnes of propellant, and a dry mass of 22.2 tonnes, meaning that they are using 17.97 tonnes of structural/tankage mass for those 411 tonnes of prop. That's a structural ratio of 22.9:1 and represents a 4.12% structural mass fraction. If you used the RD-180 with the Falcon 9's structural ratio, you'd have a 4.09% structural mass fraction. Results:

• Kerolox SSTO (altitude-compensated RD-180): Payload is 0.55% of GLOW and 9.6% of dry mass.
• Kerolox SSTO (altitude-compensated Merlin 1D): Payload is 1.41% of GLOW and 28.8% of dry mass.

And, remember: this assumes altitude compensation with NO mass penalty, which is frankly unrealistic.

I wish we had a good estimate for the TWR of the Raptor. Note that with the same structural design as the Falcon 9, an equivalent methalox launcher would have a structural ratio of 18.17:1, as the bulk density of methalox is 21% lower.

[sevenperforce]

As a basic comparison for a methalox SSTO...the Raptor is expected to have the highest-ever TWR of any liquid rocket engine. If we speculatively place this at 200:1, what does a methalox SSTO look like?

Per my table, the methalox fuel fraction is 92.3%. To get a vehicle TWR of 1.3, we need 0.65% of GLOW to be yummy Raptor goodness. Using the same structural fraction as the Falcon 9 gives 5.08% or perhaps a little lower for the tanks and body. This gets us a payload which is 1.97% of GLOW and 34.4% of dry mass...a marked improvement over a Merlin-derived SSTO, but still not enough to enable reuse.

For example, consider a speculative single-Raptor SSTO with mass-penalty-free altitude compensation. The Raptor pushes 3,050 kN of thrust at sea level; at a vehicle TWR of 1.3 that's a GLOW of 239 tonnes. Using my numbers above, that would be just 4.7 tonnes of payload to LEO, with no margin for recovery at all.