SpaceX Raptor engine - General Thread 4

[FutureSpaceTourist]

https://twitter.com/ringwatchers/status/1703798585048105339

Raptor Diagram #4 | 18th September, 2023 | "Now with high res online!"

https://ringwatchers.com/diagrams/raptor-diagrams/4

[RedLineTrain]

https://www.nextbigfuture.com/2023/09/beyond-the-spacex-raptor-engine-is-the-breakthrough-spacex-leet-1337-engine.html

Forgive my posting if this has been posted (I’m working on a phone). 

This all sounds fantastical and maybe reaching beyond the possible.

It looks like his source is the Musk biography. Has anyone got a copy of that yet?

I just listened to that part yesterday.  Raptor wasn't hitting its cost and manufacturability targets.  It was hard to manufacture and cost $2 million per.  They were only building one every three days or so.  Musk ordered a surge on Raptor to get it to $200k per, but Musk felt the effort was stale.  So he began a clean sheet design called 1337, or LEET. No material, requirement, or method was sacred on 1337.  Eventually, they paused 1337 development to go back to Raptor.  As of late last year, they were able to produce more than one Raptor a day, so obviously they made some good progress.

It was left to the reader's imagination whether the 1337 effort was Musk's way to kickstart the development team's creativity on Raptor or whether eventually the team's attention will be turned back to 1337 after the successful ramp of Raptor.

My guess is that 1337 development was only paused (not canceled) and it will be or already has been restarted.  It's also possible that the reinvigorated Raptor development is showing better than hoped progress, making 1337 development moot.

[drzerg]

High combustion chamber pressure rockets are ridiculously efficient (well above 90%) at converting combustor heat energy into exhaust velocity, and the exhausts are cold in their reference frame - to the point that water will start to condense and freeze at near to nozzle exit of high expansion ratio vac engines.

You are wrong. Water freeze in that famous rl 10 video because nozzle is cooled by liquid hydrogen. Exhaust is still very hot but it lacks radiative heat to melt that ice.

[OTV Booster]

High combustion chamber pressure rockets are ridiculously efficient (well above 90%) at converting combustor heat energy into exhaust velocity, and the exhausts are cold in their reference frame - to the point that water will start to condense and freeze at near to nozzle exit of high expansion ratio vac engines.

Raptor consumes 140kg of methane per second with a specific energy of 50MJ/kg, which is 7GW.

650 kg/sec at 3677 m/sec (Raptor-Vac) is 4.4GW of kinetic energy

An efficiency of 63%.

Which leaves you with 2.6GW of thermal energy contained in the exhaust molecules H₂O and CO₂.

Getting back to "what happens to that energy in a vacuum" question I asked, the thermal energy slowly dissipates via Stefan Boltzmann radiation, no energy is dispersed via e.g. Van der Waals forces because the molecules are too far apart, and the velocity doesn't slow down until something is hit.

So like the solar wind, it gets "thinner" farther away (the flux decreasing as function of the square of the distance), but retains all kinetic and a good portion of its thermal energy.

(this has implications for landing on the moon, for hot-staging, and for the OLM pad, but I'm just keeping it to Raptor-specific issues here on this thread)

Is that the correct answer to my question "what happens to the exhaust energy of a Raptor in a vacuum"?

Got a question not directly related to your point. How much energy does it take to move the required volume of propellant at nosebleed pressures and how efficient are the preburners?


IIUC, gas generators cycles cost ~10%, so maybe not that big a piece of the 63%.

[InterestedEngineer]

High combustion chamber pressure rockets are ridiculously efficient (well above 90%) at converting combustor heat energy into exhaust velocity, and the exhausts are cold in their reference frame - to the point that water will start to condense and freeze at near to nozzle exit of high expansion ratio vac engines.

Raptor consumes 140kg of methane per second with a specific energy of 50MJ/kg, which is 7GW.

650 kg/sec at 3677 m/sec (Raptor-Vac) is 4.4GW of kinetic energy

An efficiency of 63%.

Which leaves you with 2.6GW of thermal energy contained in the exhaust molecules H₂O and CO₂.

Getting back to "what happens to that energy in a vacuum" question I asked, the thermal energy slowly dissipates via Stefan Boltzmann radiation, no energy is dispersed via e.g. Van der Waals forces because the molecules are too far apart, and the velocity doesn't slow down until something is hit.

So like the solar wind, it gets "thinner" farther away (the flux decreasing as function of the square of the distance), but retains all kinetic and a good portion of its thermal energy.

(this has implications for landing on the moon, for hot-staging, and for the OLM pad, but I'm just keeping it to Raptor-specific issues here on this thread)

Is that the correct answer to my question "what happens to the exhaust energy of a Raptor in a vacuum"?

Got a question not directly related to your point. How much energy does it take to move the required volume of propellant at nosebleed pressures and how efficient are the preburners?


IIUC, gas generators cycles cost ~10%, so maybe not that big a piece of the 63%.

Full flow staged combustion - the work to pressurize the propellant is not wasted out the side like a gas generator. 

I'd heard that it's < 1% energy potential of the methane wasted in the turbopumps.

63% is pretty amazingly good for a heat engine.

[Hooperball]

M1-D turbine exhaust gasses, produce an additional (approximate) 4,750 lbf thrust per engine, so it's not all wasted. But just like the fuel rich combustion products from the Merlin turbines, each preburner on the R-2+ isn't running at a stoichiometric combustion ratio either. Asking for "preburner combustion efficiency" is a function of fuel/lox mass flow. Deriving impeller HP's (or turbine efficiency's) is an interesting topic because the combustion chamber basically is an injector nozzle, and if the total sum of all combustion products is 98% stoichiometric, it's hard to say what the P/T losses between turbine exhausts, injector nozzles, and chamber pressures are. Someone here knows the answer.

Changing the same topic, S25 engines still have hydraulically actuated gimballing control arms and haven't received electro-mechanical actuators like S26. Did they de-stack S-25 for an upgrade? And, do the new electro-mechanical actuators function as a screw jack with rotating stepper motor armatures outside of a threaded shaft?

S
           
                                                                 

[Tangilinear Interjar]

There's a huge amount of ship systems differences between the hydraulic and electric gimballing engines so I can confidently say they aren't making that change to S25.

There's good pictures around of the actuators but as they are completely enclosed it's hard to tell exactly what type of mechanism it uses but my money is on a recirculating ball actuator.

[OTV Booster]

High combustion chamber pressure rockets are ridiculously efficient (well above 90%) at converting combustor heat energy into exhaust velocity, and the exhausts are cold in their reference frame - to the point that water will start to condense and freeze at near to nozzle exit of high expansion ratio vac engines.

Raptor consumes 140kg of methane per second with a specific energy of 50MJ/kg, which is 7GW.

650 kg/sec at 3677 m/sec (Raptor-Vac) is 4.4GW of kinetic energy

An efficiency of 63%.

Which leaves you with 2.6GW of thermal energy contained in the exhaust molecules H₂O and CO₂.

Getting back to "what happens to that energy in a vacuum" question I asked, the thermal energy slowly dissipates via Stefan Boltzmann radiation, no energy is dispersed via e.g. Van der Waals forces because the molecules are too far apart, and the velocity doesn't slow down until something is hit.

So like the solar wind, it gets "thinner" farther away (the flux decreasing as function of the square of the distance), but retains all kinetic and a good portion of its thermal energy.

(this has implications for landing on the moon, for hot-staging, and for the OLM pad, but I'm just keeping it to Raptor-specific issues here on this thread)

Is that the correct answer to my question "what happens to the exhaust energy of a Raptor in a vacuum"?

Got a question not directly related to your point. How much energy does it take to move the required volume of propellant at nosebleed pressures and how efficient are the preburners?


IIUC, gas generators cycles cost ~10%, so maybe not that big a piece of the 63%.

Full flow staged combustion - the work to pressurize the propellant is not wasted out the side like a gas generator. 

I'd heard that it's < 1% energy potential of the methane wasted in the turbopumps.

63% is pretty amazingly good for a heat engine.

Yeah, 63% is awesome. Still, the question of where it goes is interesting. The only other losses I came come up with are that long cooling circuit on the methane side, generating ullage gas and unrealized potential energy from under expansion.


Edit: almost forgot, fuel rich, film cooling and the quench from the regen cooling cause chemical species that don't react until past the bell, if they react at all. Maybe this all comes under the unrealized potential energy I already mentioned.

[OTV Booster]

Got another question. If the Raptor were built out of unobtanium and they could burn stoic, and ditch film cooling and the regen circuit, would this increase thrust or ISP? Or is it one of those 'it depends' things.


Ahhh, an epiphany as I wrote. Thrust is like volts and ISP is like amperage. Need them both to create a unit of work. The unobtainium engine will do more work and how it divvies thrust and ISP is a design decision. So, it depends.

[tbellman]

Got another question. If the Raptor were built out of unobtanium and they could burn stoic, and ditch film cooling and the regen circuit, would this increase thrust or ISP? Or is it one of those 'it depends' things.

It's one of those "it's (slightly) complicated" things. 

Perfect stoichiometric combustion gives you the most energy per kilogram of fuel+oxidizer.  However:

A) Perfect stoichiometric combustion rarely exists in the real world.  There will always be some oxidizer molecules and some fuel molecules that fail to meet up and combust.  In practice, having a slightly fuel rich ratio, tends to give you a little bit more energy, as it is then more likely that all the oxidizer is combusted.  While you will then get more fuel that is not combusted, the fuel is usually lighter than the oxidizer.

B) Maximum energy is usually not what gives you most thrust or highest specific impulse.

Thrust in a rocket engine comes from the ideal gas law, pV = nRT; pressure (p) times volume (V) equals amount of substance in the exhaust (n) times the gas constant (R, equivalent to the Boltzmann constant) times temperature (T).  You want to maximize the left hand side, which of course means maximizing the right hand side.  The gas constant is a natural constant, which we can't change; and it is the same for all gases.

The temperature depends of course on the heat energy released by the combustion process.  But also on the composition of the exhaust gases.  Different gases require different amounts of energy to raise the temperature the same amount; the specific heat capacity.  Water, e.g, is fairly high, at 4.2 kJ per gram and kelvin, while carbon dioxide is more like 1.2 kJ/g·K, and methane is 2.2 kJ/g·K; and hydrogen is on the order of 16 kJ/g·K.  (Note also that those figures vary with temperature.)  Thus, having more methane than water in the exhaust, can actually increase the temperature for the same amount of released combustion heat.

But more important is the amount of substance in the exhaust.  Because that is not measured in kilograms, as one could naïvely expect, but in number of particles (molecules and/or atoms).  To maximize that per kilogram of fuel+oxidizer (and the mass of fuel+oxidizer is equal to the mass of the exhaust in a closed cycle rocket engine), you want as light molecules as possible in the exhaust.  Carbon dioxide has a molecular mass of 44, and water has a molecular mass of 18, while methane is down at 16.  And even better, methane can dissociate at those temperatures, i.e. split into its component atoms, or at least split off one or more hydrogen atoms.  Dissociation increases the number of particles, increasing n.

Thus, running a methane+oxygen rocket engine slightly fuel rich actually increases the specific impulse.  I think the maximum is around a 1:3.6 ratio between methane and oxygen measured by mass (i.e. 22% methane), while the stoichiometric ratio is almost exactly 1:4.0 (i.e. 20% methane).

A hydrogen+oxygen engine is even more skewed; stoichiometric ratio is 1:8.0 (11% hydrogen), but you get the highest specific impulse at around 1:4 (20% hydrogen).

And the real world is even more complicated, because at the pressures in a rocket engine, the exhast is not a perfectly ideal gas.  The ideal gas law is a simplified model of the actual real world.  We have also ignored the fact that real world combustion of methane and oxygen doesn't just produce water and carbon dioxide, but also small amounts of several other combinations of carbon (C), hydrogen (H) and oxygen (O), e.g. carbon monoxide (CO), hydroxyl (OH), and many more.  But the above should hopefully give an idea of the basics.

[FutureSpaceTourist]

https://twitter.com/nasaspaceflight/status/1704710878652694868

SpaceX put an RVac on a float and took it on a parade at McGregor:

"SpaceX, known for its secrecy, hasn’t had a public showing like this before."

A bit of video and some pictures in the article:

https://www.kwtx.com/2023/09/20/spacex-participates-mcgregor-founders-day-parade-stuns-onlookers-with-huge-raptor-vacuum-engine/

[ETurner]

High combustion chamber pressure rockets are ridiculously efficient (well above 90%) at converting combustor heat energy into exhaust velocity, and the exhausts are cold in their reference frame - to the point that water will start to condense and freeze at near to nozzle exit of high expansion ratio vac engines.

Raptor consumes 140kg of methane per second with a specific energy of 50MJ/kg, which is 7GW.

650 kg/sec at 3677 m/sec (Raptor-Vac) is 4.4GW of kinetic energy

An efficiency of 63%.

Which leaves you with 2.6GW of thermal energy contained in the exhaust molecules H₂O and CO₂.

Getting back to "what happens to that energy in a vacuum" question I asked, the thermal energy slowly dissipates via Stefan Boltzmann radiation, no energy is dispersed via e.g. Van der Waals forces because the molecules are too far apart, and the velocity doesn't slow down until something is hit.

So like the solar wind, it gets "thinner" farther away (the flux decreasing as function of the square of the distance), but retains all kinetic and a good portion of its thermal energy.

(this has implications for landing on the moon, for hot-staging, and for the OLM pad, but I'm just keeping it to Raptor-specific issues here on this thread)

Is that the correct answer to my question "what happens to the exhaust energy of a Raptor in a vacuum"?

Got a question not directly related to your point. How much energy does it take to move the required volume of propellant at nosebleed pressures and how efficient are the preburners?


IIUC, gas generators cycles cost ~10%, so maybe not that big a piece of the 63%.

Full flow staged combustion - the work to pressurize the propellant is not wasted out the side like a gas generator. 

I'd heard that it's < 1% energy potential of the methane wasted in the turbopumps.

63% is pretty amazingly good for a heat engine.

Yes, but a rocket engine isn't a heat engine.

Even Wikipedia is confused about heat engines and rocket efficiency:

Combustion is most frequently used for practical rockets, as the laws of thermodynamics (specifically Carnot's theorem) dictate that...

This makes no sense. Carnot efficiency = 1 – T_low /T_high, but there is no T_low, because there is no heat sink! "Heat engine" is a technical term that doesn't just mean "an engine that uses thermal energy", and a rocket engine isn't a heat engine.

In the limit of large expansion ratio (and a few idealizations, like fast equilibration of thermal velocity, vibration, and rotation), a rocket engine should convert thermal kinetic energy in the combustion chamber to directed kinetic energy in the exhaust with close to 100% efficiency. I question the 63% number. How does the combustion energy density of the propellant (not the methane) compare to the exhaust kinetic energy?

A fully expanded rocket plume gets cold at high altitudes, so there's not much thermal energy left:

...the condensation of water vapor and carbon dioxide can start at distances of 120–170 m and 450–650 m from the engine nozzle, respectively.

Condensation of water vapor and carbon dioxide in the jet exhausts of rocket engines...

[edit:]

For nozzles that are used in vacuum or at very high altitude, it is impossible to match ambient pressure; rather, nozzles with larger area ratio are usually more efficient [...] as the temperature of the gas in the nozzle decreases, some components of the exhaust gases (such as water vapour from the combustion process) may condense or even freeze.

(Rocket_engine_nozzle#Vacuum_use)

[scaesare]

There's a huge amount of ship systems differences between the hydraulic and electric gimballing engines so I can confidently say they aren't making that change to S25.

There's good pictures around of the actuators but as they are completely enclosed it's hard to tell exactly what type of mechanism it uses but my money is on a recirculating ball actuator.

I wonder if they are a scaled up unit similar to some they have developed for Optimus:

[FutureSpaceTourist]

https://twitter.com/considercosmos/status/1709384918407487598

Cybertruck towing a raptor vac at Starbase!

[oldAtlas_Eguy]

Yes but the interesting thing about that Raptor is its number #305!!!!

The Raptor tracking pages needs a new page: 100s, 200s, and now 300s.

[wannamoonbase]

Yes but the interesting thing about that Raptor is its number #305!!!!

The Raptor tracking pages needs a new page: 100s, 200s, and now 300s.

The volume of Raptors is the most amazing part of raptor development.

They've only mounted and used about a third of them at this point.  So plenty of scrap of older ones lay around, but they have served their purpose to develop the engine and manufacturing.

[Tangilinear Interjar]

I wonder if they make design changes to the Raptors and skip numbers kind of like they do with Starship numbers. For example they are at 300 something but have only fully assembled 200 ish?

[Asteroza]

I wonder if they make design changes to the Raptors and skip numbers kind of like they do with Starship numbers. For example they are at 300 something but have only fully assembled 200 ish?

First digit is a design series marker, so 3xx is equivalent to mark 3?

[eriblo]

I wonder if they make design changes to the Raptors and skip numbers kind of like they do with Starship numbers. For example they are at 300 something but have only fully assembled 200 ish?

First digit is a design series marker, so 3xx is equivalent to mark 3?

I have not seen anything to indicate that the first digit is anything else than the 10² digit in the three digit sequential Raptor 2 serial number.

[SkyRate]

I wonder if they make design changes to the Raptors and skip numbers kind of like they do with Starship numbers. For example they are at 300 something but have only fully assembled 200 ish?

Pushing the boundaries as far as they are doing, we can be sure they have burned through (pun intended) a lot of engines in tests.