Author Topic: SpaceX Raptor engine - General Thread 4  (Read 1204745 times)

Offline rsdavis9

Re: SpaceX Raptor engine - General Thread 4
« Reply #2640 on: 11/30/2023 01:39 pm »
Someplace there is a graph of O/F and isp. And I think 3.6 is already optimum for isp.
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Online eriblo

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Re: SpaceX Raptor engine - General Thread 4
« Reply #2641 on: 11/30/2023 01:57 pm »
Just handwavium here.

As the mixture is changed, the combustion temperature will change.

Lots of factors involved.

What I'm thinking is that in certain circumstances, total thrust is really important.  That would be early on when velocity is low and gravity losses are high.

But, as orbital velocity is approached, gravity losses drop to zero, and thrust becomes unimportant, or far less important.

Under those circumstances, is it possible to adjust the fuel oxidizer ratio of the raptor to trade off thrust for ISP?

That is what I'm asking.
We do not know to what degree Raptor can adjust the mixture ratio while running. Just because you have separate turbines and pumps does not mean that it is easy, they are still interconnected at the preburners and the injectors must be able to handle the change. Not having ignitors in the main chamber might also play a role.

Some engines can and do adjust mixture ratio in flight, but I do not know if it is used to optimize Isp over the trajectory as opposed to minimizing propellant residuals at the end of the burn.

Online Robotbeat

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Re: SpaceX Raptor engine - General Thread 4
« Reply #2642 on: 11/30/2023 02:00 pm »
Someplace there is a graph of O/F and isp. And I think 3.6 is already optimum for isp.
Depends on chamber pressure and exit pressure. Near-stoichiometric, the problem is that your engine melts, not necessarily that the Isp drops. Also, the more oxygen, the greater your bulk density and so the lighter the tanks & pumps are for a given propellant mass.
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Offline tbellman

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Re: SpaceX Raptor engine - General Thread 4
« Reply #2643 on: 11/30/2023 02:02 pm »
So,

I've seen the ratio of 4 floated around as stochiometric.

but, I would think that.

CH4 + 2 O2 -> CO2 + 2 H2O 

or am I missing something?  That would be number of molecules.

But based on weight.

2 O2 weighs 64 and CH4 weighs 16, so based on weight, that would be where the 4 to 1 mixture comes from.

So, 3.6 would be fuel rich.

In the field of rocketry, ratios between oxidizer and fuel are almost always specified as the mass ratio.  So when Starship has an oxygen-methane ratio of 3.6:1, what is meant is that it uses 3.6 tonnes of oxygen for each tonne of methane.  The Space Shuttle had an oxygen-hydrogen ratio of about 6:1, again by mass.  And the stoichiometric ratio under perfect combustion is almost exactly 4:1 (63.998:16.043 with a few more decimals) by mass.  So both Starship and Space Shuttle uses/used fuel-rich combustion.

Chemists on the other hand usually measures reaction ratios in moles.  So in the field of chemistry, the stoichiometric combustion ratio of oxygen and methane would be specified as 2:1.  (I think chemists usually specify the fuel first, so they would then rather say the stoichiometric ratio is 1:2.)

Rocketry uses mass, because that's the property that is most useful in this field.  Mass is e.g. relevant in Tsiolkovsky's rocket equation, and in many other calculations.  Mass is also easy to measure; just put your fuel and oxidizer tanks on a pair of scales.  The number of molecules (or moles) basically never show up in rocketry calculations.

Occasionally, volumes are also used.  But you should expect it to be called out explicitly when done.



An oxygen-methane ratio of 3.6:1 by mass is as said already fuel-rich, and is pretty close to the optimum for specific impulse.

For oxygen-hydrogen, then 8:1 is the stoichometric ratio by mass, but the optimum for specific impulse is around 4:1, i.e. twice as much hydrogen as in the stoichiometric case.  Most hydrolox rockets however use a somewhat less fuel-rich ratio, even though it gives worse specific impulse.  They do so because of hydrogens very low density.  By going to around 6:1 oxygen-hydrogen, you need less volume in your tanks, and the dry mass of the tanks thus decrease.  It is also easier to get higher absolute thrust with a denser mixture (fuel pumps can be smaller and use less power, fuel lines, combustion chamber, engine bell, et.c., can be smaller).

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Re: SpaceX Raptor engine - General Thread 4
« Reply #2644 on: 11/30/2023 02:09 pm »
I meanÖ hypothetically, with a high enough chamber pressure and a long enough nozzle and with unobtanium chamber wall materials, even hydrolox benefits from near-stoichiometric as the actual final velocity of the exhaust depends on the chemical input energy, which is maximized by being near stoichiometric.

 BUT in practice you end up limited by TEMPERATURE, so in that case you want to minimize molecular mass of the exhaust while riding the temperature limit. Also, the long nozzle ends up being heavy, too, so gamma matters, too.
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Offline edzieba

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Re: SpaceX Raptor engine - General Thread 4
« Reply #2645 on: 11/30/2023 02:32 pm »
I meanÖ hypothetically, with a high enough chamber pressure and a long enough nozzle and with unobtanium chamber wall materials, even hydrolox benefits from near-stoichiometric as the actual final velocity of the exhaust depends on the chemical input energy, which is maximized by being near stoichiometric.

 BUT in practice you end up limited by TEMPERATURE, so in that case you want to minimize molecular mass of the exhaust while riding the temperature limit. Also, the long nozzle ends up being heavy, too, so gamma matters, too.
Even with the infinite filigree spherical-cow nozzle, I don't think stoich is the optimum. If your optimisation target is maximising ISP, then that's the same as maximising exhaust velocity. If we were using an NTR with a single molecule as propellant, then core/exhaust temperature would correlate directly to ISP (same molecule but hotter = same molecule moving faster. But for chemical combustion, changing mixture ratio does not just change combustion temperature, but also the resulting reaction products. If you move away from stoich your reaction products may start to become cooler, but if that also means your reaction products are also lighter then they will overall be faster too (so higher ISP).
Different propellant combinations will trade this differently, but since rocket props generally tend to have a fuel molecule that is lighter than its oxidiser molecule (in bulk) it is usually the case that running slightly rich will be produce a lighter exhaust and hence a faster exhaust than perfectly stoich. There may be some weird combinations whose optimum mixture ratio (for ISP) is lean, but a hot oxidising combustion chamber seems like a bad time for everyone involves, not least of which the poor combustion chamber.

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Re: SpaceX Raptor engine - General Thread 4
« Reply #2646 on: 11/30/2023 03:02 pm »
It is worth noting that it is not the molecular weight of the oxidizer/fuel that is important (although there are secondary effects with regard to bulk density) but rather the molecular weight of the products.

H, H2, OH and CO are better than CO2 and H2O (with a mixture of H and C you are mostly trading for CO). Simpler molecules also have less vibrational and rotational modes so more of the total energy is in translation (i.e. velocity).

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Re: SpaceX Raptor engine - General Thread 4
« Reply #2647 on: 11/30/2023 03:18 pm »
Conservation of energy. Your final exhaust velocity cannot, in a chemical rocket engine, exceed the available chemical energy, or rather:

chemical specific energy >= specific kinetic energy = (1/2)*v^2

The efficiency of converting that chemical energy to kinetic energy depends on the details of the rocket engine, but starting with very high chamber pressure and with the engine in a vacuum, a long enough nozzle (insulated, letís say), the efficiency can approach 100%, in which case the details like exhaust molecular mass donít really matter.
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Offline BarryKirk

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Re: SpaceX Raptor engine - General Thread 4
« Reply #2648 on: 12/01/2023 11:06 am »
Conservation of energy. Your final exhaust velocity cannot, in a chemical rocket engine, exceed the available chemical energy, or rather:

chemical specific energy >= specific kinetic energy = (1/2)*v^2

The efficiency of converting that chemical energy to kinetic energy depends on the details of the rocket engine, but starting with very high chamber pressure and with the engine in a vacuum, a long enough nozzle (insulated, letís say), the efficiency can approach 100%, in which case the details like exhaust molecular mass donít really matter.

That makes perfect sense.  One would think that the maximum chemical specific energy would occur at stochiometric ratio.  Of course, that may happen at temperatures that would melt the engine.

Offline edzieba

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Re: SpaceX Raptor engine - General Thread 4
« Reply #2649 on: 12/01/2023 11:19 am »
Conservation of energy. Your final exhaust velocity cannot, in a chemical rocket engine, exceed the available chemical energy, or rather:

chemical specific energy >= specific kinetic energy = (1/2)*v^2

The efficiency of converting that chemical energy to kinetic energy depends on the details of the rocket engine, but starting with very high chamber pressure and with the engine in a vacuum, a long enough nozzle (insulated, letís say), the efficiency can approach 100%, in which case the details like exhaust molecular mass donít really matter.
True, but unhelpful. The optimisation target (for maximising  Specific Impulse) is not necessarily maximum combustion efficiency, for the reasons mentioned earlier. Maximum ISP is not always your optimisation target either, e.g. for first stages where thrust density or impulse density may be preferable. Though even in those cases, shoving extra mass through the engine that does not completely combust also leads you away from stoich.

Offline rsdavis9

Re: SpaceX Raptor engine - General Thread 4
« Reply #2650 on: 12/01/2023 11:59 am »
Conservation of energy. Your final exhaust velocity cannot, in a chemical rocket engine, exceed the available chemical energy, or rather:

chemical specific energy >= specific kinetic energy = (1/2)*v^2

The efficiency of converting that chemical energy to kinetic energy depends on the details of the rocket engine, but starting with very high chamber pressure and with the engine in a vacuum, a long enough nozzle (insulated, letís say), the efficiency can approach 100%, in which case the details like exhaust molecular mass donít really matter.

This is one of the unintuitive things about chemical propulsion. The molecular mass of the combustion products makes a difference.
In a metholox:
CO2 - 44
H2O - 18
If run fuel rich
CO - 28

So you get a higher exhaust velocity with lower molecular mass and higher exhaust velocity is higher isp.

EDIT: Hmm but you are right.
I tried 1000 for the expansion ration and changed from O/F 3.6 to 4.0 and got
3.6 -> 399.96s
4.0 -> 400.34s
« Last Edit: 12/01/2023 12:11 pm by rsdavis9 »
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Offline InterestedEngineer

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Re: SpaceX Raptor engine - General Thread 4
« Reply #2651 on: 12/01/2023 03:06 pm »
Conservation of energy. Your final exhaust velocity cannot, in a chemical rocket engine, exceed the available chemical energy, or rather:

chemical specific energy >= specific kinetic energy = (1/2)*v^2

The efficiency of converting that chemical energy to kinetic energy depends on the details of the rocket engine, but starting with very high chamber pressure and with the engine in a vacuum, a long enough nozzle (insulated, letís say), the efficiency can approach 100%, in which case the details like exhaust molecular mass donít really matter.

This is one of the unintuitive things about chemical propulsion. The molecular mass of the combustion products makes a difference.
In a metholox:
CO2 - 44
H2O - 18
If run fuel rich
CO - 28

So you get a higher exhaust velocity with lower molecular mass and higher exhaust velocity is higher isp.

EDIT: Hmm but you are right.
I tried 1000 for the expansion ration and changed from O/F 3.6 to 4.0 and got
3.6 -> 399.96s
4.0 -> 400.34s

Here's the equation that uses the molecular weight of exhaust.  In it you can see the tradeoffs between chamber temperature, pressure,  molecular weight, etc.

Equation is just as important as the Rocket Equation

https://en.wikipedia.org/wiki/De_Laval_nozzle#Exhaust_gas_velocity

Online LouScheffer

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Re: SpaceX Raptor engine - General Thread 4
« Reply #2652 on: 12/01/2023 05:43 pm »
Since, the Raptor has separate turbo pumps for fuel and oxidizer, how easy or hard would it be to burn in a non stoichiometric ratio?

I'm thinking that by going fuel rich, the molecular weight of the exhaust products could be reduced and thereby the ISP could be increased.

So, when a star ship is approaching orbital velocity would it make sense to change the fuel/oxidizer ratio to increase the ISP at the cost of thrust.

Would it make sense for a fuel tanker to be fuel rich, since gravity losses don't occur when doing orbital transfers, and there is very little gravity losses for lunar landing/takeoff.
This was standard procedure during Apollo.  From Schemes for Enhancing the Saturn V Translunar Payload Capability:
Quote
Early in that rocketís flight, we set the burning-mixture ratio at 5.5 to 1 (5.5 pounds of oxidizer for every pound of fuel). But 70 percent of the way through the burn we abruptly shifted that mixture ratio to a lower value of 4.5 to 1. As the small graphs in Figure 1 indicate, this shift in the mixture ratio provided the rocket with high thrust early in its flight at a slightly lower specific impulse.* Then, following the Programmed Mixture Ratio shift, it had a lower thrust, but a higher specific impulse.

As Raptor has separate pumps as you point out, they should have no trouble doing this optimization (or it's continuous adjustment cousin) should they decide this is helpful.

Offline InterestedEngineer

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Re: SpaceX Raptor engine - General Thread 4
« Reply #2653 on: 12/01/2023 06:23 pm »
Since, the Raptor has separate turbo pumps for fuel and oxidizer, how easy or hard would it be to burn in a non stoichiometric ratio?

I'm thinking that by going fuel rich, the molecular weight of the exhaust products could be reduced and thereby the ISP could be increased.

So, when a star ship is approaching orbital velocity would it make sense to change the fuel/oxidizer ratio to increase the ISP at the cost of thrust.

Would it make sense for a fuel tanker to be fuel rich, since gravity losses don't occur when doing orbital transfers, and there is very little gravity losses for lunar landing/takeoff.
This was standard procedure during Apollo.  From Schemes for Enhancing the Saturn V Translunar Payload Capability:
Quote
Early in that rocketís flight, we set the burning-mixture ratio at 5.5 to 1 (5.5 pounds of oxidizer for every pound of fuel). But 70 percent of the way through the burn we abruptly shifted that mixture ratio to a lower value of 4.5 to 1. As the small graphs in Figure 1 indicate, this shift in the mixture ratio provided the rocket with high thrust early in its flight at a slightly lower specific impulse.* Then, following the Programmed Mixture Ratio shift, it had a lower thrust, but a higher specific impulse.

As Raptor has separate pumps as you point out, they should have no trouble doing this optimization (or it's continuous adjustment cousin) should they decide this is helpful.

There are McGregor videos showing them varying the mix.

I've seen the videos, but can't remember which ones.

Offline livingjw

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Re: SpaceX Raptor engine - General Thread 4
« Reply #2654 on: 12/01/2023 10:58 pm »
Someplace there is a graph of O/F and isp. And I think 3.6 is already optimum for isp.

Here you go. If you lifted off using MR for maximum density impulse (3.75) and ended the flight at maximum Isp MR (3.45), and you sized your  tanks accordingly, you would increase your delta V a little.

John
« Last Edit: 12/01/2023 11:05 pm by livingjw »

Offline Slarty1080

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Re: SpaceX Raptor engine - General Thread 4
« Reply #2655 on: 12/07/2023 10:12 am »
Quick Raptor question for a calculation, does anyone know what the diameter is of the vacuum Raptor nozzle exit?
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Offline Brigantine

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Re: SpaceX Raptor engine - General Thread 4
« Reply #2656 on: 12/07/2023 12:18 pm »
what the diameter is of the vacuum Raptor nozzle exit?
I had 2.4m

source: "The vacuum flight version, with a nozzle exit diameter of 2.4 m (7.9 ft)"

But the cite note is from 2018, so maybe I'm using the wrong out of date value

Offline rsdavis9

Re: SpaceX Raptor engine - General Thread 4
« Reply #2657 on: 12/07/2023 12:25 pm »
what the diameter is of the vacuum Raptor nozzle exit?
I had 2.4m

source: "The vacuum flight version, with a nozzle exit diameter of 2.4 m (7.9 ft)"

But the cite note is from 2018, so maybe I'm using the wrong out of date value

I have 2.32m from John L.
Supposedly pixel counted from mcgregor.
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Re: SpaceX Raptor engine - General Thread 4
« Reply #2658 on: 12/13/2023 09:41 pm »
https://twitter.com/groundtruthpics/status/1735066627165561320

Quote
A wild Raptor Engine #221 on December 13, 2023 #Starbase
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Offline tenkendojo

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Re: SpaceX Raptor engine - General Thread 4
« Reply #2659 on: 12/15/2023 04:30 pm »
https://twitter.com/groundtruthpics/status/1735066627165561320

Quote
A wild Raptor Engine #221 on December 13, 2023 #Starbase
.
📸Me for WAI Media @FelixSchlang

At first I thought they have truncated the engine bell, but after looking at it more closely it appears that they just moved the shield (the black ring around the bell) further down, creating an optical illusion of a shorter bell. I think the nozzle shape remained the same.

 

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