Raptor, Methane and FFSC

[MathiLpHD]

Hello!

I have some questions about FFSC, Methane and their combination in the Raptor engine (and it's vacuum version) SpaceX develops right now.

First some links of my sources:

http://www.scienceforums.net/topic/81051-staged-combustion-rocket-engines/

http://space.stackexchange.com/questions/3161/why-is-spacex-considering-methane-as-fuel-for-their-next-engine-the-raptor

https://en.wikipedia.org/wiki/Staged_combustion_cycle#Full-flow_staged_combustion_cycle

https://en.wikipedia.org/wiki/Comparison_of_orbital_rocket_engines

https://en.wikipedia.org/wiki/Liquid_rocket_propellant#Bipropellants

So, methane gains about 10s of specific impuls over rp-1 and has just about 20% less bulk density. If you have a look at the comparison of orbital rocket engines and sort it by highest pressure and have a look on the RD-180 1st stage engine: It has a atm isp of ~310s and a vacuum isp of ~340s but "just" uses  oxygen-riched staged combustion with RP-1/LOX. So the isp of the Raptor engine should gain 10s from the higher isp of the fuel and 10s for using FFSC compared to the RD-180. So, if SpaceX optimizes the 1st stage version of the Raptor engine, they should get an isp of ~330s atm and ~360s vac, maybe even 370s. So i think the official isp of 321s atm and 363s vac are estimations for the first version or maybe for their tests and will be increased in the later versions.

But as far as i know there is no official isp value for the upper stage version of the Raptor engine. So, let's have a look at the RD-0124. It uses the same combustion cycle as the RD-180 but has nearly half the camber pressure, a vacuum nozzle and uses RP-1/LOX. It has a vac isp of ~360s. I don't know why they used such low chamber pressure, maybe because they wanted to reduce weight. So with the full chamber pressure of oxygen-riched staged combustion it should get an vac isp of ~380s (compare Merlin 1D (280s/310s) with RD180 (310s/340s) and Merlin 1D Vacuum (~350s) with RD-0124 (~360s) --> +20s with full pressure). So, if they do FFSC with full pressure they should get around 400s vac isp for the upper stage Raptor engine. If they could archieve that, this would be the first chemical non-LH2 engine that reaches the 400s isp mark and nearly getting the same isp of some early LH2 engines.

So my question: Is that a realistic calculation i did there? Or won't they use the full pressure of an FFSC (~31MPa)? And could you please list all advantages/disadvantages of Methane and FFSC compared to other propellants/engine cycles?

[Elthy]

You cant just add stuff like this. There is a theoretical limit for ISP for each fuel/oxidizer combination, i doubt CH4/LOX reaches 400s but im not able to find the numbers right now...

[MathiLpHD]

@Elthy

Why not? Chamber pressure increases the isp of every propellant. There is not a real isp maximum for fuels. Maybe a limit for the engineers because they can't increase the chamber pressure anymore, but not a theoretical limit.

[cfds]

There is a theoretical limit, the kinetic energy of the ejected propellants can not be higher than the energy released by burning the fuel.

[MathiLpHD]

Well, that's true but that doesn't limit the isp. The isp is just limited by the exhaust velocity. But you can always increase the exhaust velocity by decreasing the diameter of your nozzle/pumping more fuel into the burning chamber. The fuel flow must be the same for the burning chamber and the nozzle, so the speed of the exhaust will increase if you decrease the diameter of the nozzle or pump more fuel into your burning chamber. So the only real limits are the material limits of the engine. You can't increase the pressure through the nozzle to thousends of bars. The nozzle wouldn't be able to handle that. But in theory, there is no limit for the isp. But that's also dependet on how you define the maximum isp. Because the fuel pumps can't produce a unlimited fuel flow. At least if you power them with the fuel it's pumping. If you power them with an electric engine, you should be able to get a realy high isp, but the battery you would have to use would be to heavy to use in a rocket. And then you would use an additional energy source which maybe conflicts with the definition of the maximum isp. So there is a theoretical isp maximum for every type of engine with it's fuel, but there isn't a general isp maximum for a fuel because it depends on the type of engine cycle you use.

Edited March 3, 2016 by MathiLpHD

[shynung]

But I_sp is just exhaust velocity times standard gravity. Whatever you do that increases the exhaust velocity automatically increases the I_sp.

On chemical rockets, exhaust velocity is limited by type of propellant used. For a given chemical propellant combo, only a certain amount of energy is released on reaction(example: at a stoichiometric fuel/oxidizer ratio, LH2 burned with LOX generates some 286 kJ/mol of hydrogen, 2H₂ + O₂ = 2H₂O + 572 kJ). The maximum exhaust velocity possible from each combination is known by assuming that all of the energy released during the reaction goes to accelerating the reaction products. Since the energy released are tied to the propellant mass reacted, there is a limit on Isp for a given chemical propellant, no matter the engine.

Previous answer on the matter from the resident physicist:

Quote

On 11/27/2014 at 0:49 PM, shynung said:

I've seen in one mod pack a monopropellant fueled engine that manages an I_SP of about 410 seconds, which happens to have an oversized nozzle bell. My question is, how far can one increase the I_SP of an engine through throat-and-nozzle geometry alone, for a given type of fuel?

 

On 11/27/2014 at 3:59 PM, K^2 said:

Assuming opration in vacuum, inviscid flow, and massless bell, you want it running to infinity to get 100% of thermal energy turned into thrust. In practice, gains start getting insignifficant pretty fast. But this "infinite bell" approach lets you compute the absolute maximum I_SP that a particular fuel mixture can have. Simply find the amount of heat produced, and assume that every molecule of exhaust has equal energy. Then compute total impulse per mass of material. If you do this for real fuels, you'll find that this maximum is somewhat optimistic. (I believe, it works out to well over 500s for LH2/LOX. Don't feel like doing the math right now.) The main reason is that exhaust does not have time to fully thermalize between the degrees of freedom, and that's because practical limitations requrie one to cut the nozzle short. Of course, there is some turbulence and drag as well that reduce efficiency.

Operation in atmosphere is a different story. Not only extending bell past certain cutoff doesn't give you extra thrust, but it actually reduces it. Which is why engines designed for operation in atmosphere are so much shorter.

 

 

 

Edited March 3, 2016 by shynung

[MathiLpHD]

@shynung

As i said: If you add energy by powering the fuel pumps with electric energy, you should get a higher isp. So it depends on how you define the maximum isp... But i think you should be able to get an isp of ~400 out of CH4/LNG + LOX...

Update1: So i did the calculations:

CH4 + 2 O2 => 2 H20 + CO2 + ~790 kJ/Mol

1000g*790kJ/Mol/(12g/Mol+4g/Mol+64g/Mol) => 9856 kJ/kg - Energy released by 1kg CH4 + LOX mixture => 9856000 J/kg (Standard unit)

E=(m*v^2)/2 => v = sqrt(2*E/m)

v = sqrt(2*9856000J/kg) = 4440 m/s

4440/9.81 = 453s

So yes, it is possible to reach 400s with CH4.

Update2: For LH2/LOX it is ~575s.

Edited March 3, 2016 by MathiLpHD

[sgt_flyer]

21 minutes ago, MathiLpHD said:

@shynung

As i said: If you add energy by powering the fuel pumps with electric energy, you should get a higher isp. So it depends on how you define the maximum isp... But i think you should be able to get an isp of ~400 out of CH4/LNG + LOX...

the RL-10 rocket engine expander cycle turbine assembly is rated at around 500kilowatts - just for driving the pumps. :

https://books.google.fr/books?id=JnoZTbVLx0MC&pg=PA138&lpg=PA138&dq=rl10+rocket+turbine+kilowatts&source=bl&ots=C5v9UJOYMY&sig=JXtvnTtF36k8clS6XLGmOmEPBxI&hl=fr&sa=X&ved=0ahUKEwi7q8uquqTLAhVKnBoKHS6LBuUQ6AEIIzAB#v=onepage&q=rl10%20rocket%20turbine%20kilowatts&f=false

current batteries or electric generators simply have nowhere near the required power densities to match the energy rocket engines are capable of extracting from chemical reactions - not without severely increasing the dry mass of your rocket.

only on very small rockets it could moderatly make sense to have electric driven pumps from a simplicity standpoint (the engine would be way less complex to design). (and even then, it's going to be less efficient because it calls for more dry mass) ISP isn't the magic answer to everything. dry mass plays a huge part too. 

Edited March 3, 2016 by sgt_flyer

[MathiLpHD]

@sgt_flyer

Well i just said that it depends on how you define the maximum isp. I wouldn't make an engine powered with electric fuel pumps. And i calculated some maximum isps in the comment above: ~453s for CH4 + LOX and ~575s for LH2 + LOX. And as far as i heard, FFSC saves some weight because it doesn't put so much stress on the turbines...

[shynung]

@MathiLpHD

For a given thrust and propellant mix, a gas generator system is actually lighter than a comparable FFSC system. In other words, GG has a higher TWR, but the FFSC compensates this by being more efficient, i.e. higher I_sp. Here's why.

In an gas generator setup, the turbopump exhausts into either ambient atmosphere or space vacuum. Compared to a combustion chamber, these environments have very low to negligible pressure. This means the turbopump's preburner encounters very little back pressure, and the combustion conditions can be set up to be identical to that of the main combustion chamber. This would save on costs, since the preburner and the main combustion chamber can be made of the same material, simplifying logistics. Some designs cut costs even further by discarding the preburner altogether and simply pipe some of the hot gases from the main combustion chamber straight to the turbopump, simplifying construction. That's their advantage.

The main disadvantage is that this design loses efficiency because the propellant used to drive the turbine is now unavailable to be used for thrust. The exhaust of the turbine itself has very little energy left, and won't contribute much thrust.

 

In a staged combustion setup, whether full-flow or otherwise, the turbopump exhausts into the main combustion chamber. Since the main combustion chamber has a high pressure, the turbine encounters high back pressures (its exhaust gases has to push against the main chamber). This means the turbopump's preburner needs to have even higher pressures than that of the main combustion chamber, which often means the turbopump and plumbing needs to be made out of exotic materials ($$$!). Not only that, if the fuel is the kind that leaves residue on the machinery (hydrocarbon fuels, including methane, sometimes doesn't burn properly, and leaves solid monatomic carbon rather than CO₂), the resulting turbine exhaust would foul the main combustion chamber (a GG system would simply throw it aside). Also, all those plumbing are heavier than a comparable GG system. That's their disadvantage.

Their main advantage is, after all the hoops have been passed through, the engine uses all the fuel it sucks in for thrust. It's more efficient than a GG system of a comparable thrust rating.

Edited March 3, 2016 by shynung

[MathiLpHD]

@shynung

As far as i know, they can run the turbopump at cooler temperatures due to the lower temperatures of methane or so. So that they don't have to use exotic materials. And the russians already did oxygen-riched staged combustion with RP-1 which has much more carbon parts than methane, so that shouldn't be a real problem.

[sevenperforce]

To answer why you can't simply keep increasing chamber pressure to get higher ISP: eventually, as you approach that upper theoretical limit, your fuel no longer has any more chemical energy available to increase the heat and pressure in the combustion chamber. 

Full-flow staged combustion has the disadvantage of requiring two separate preburners. However, this means the requirements on each one is lower, and since you already need separate pumps, this isn't much additional weight cost. 

And yes, the Russians did oxy-rich staged combustion with RP1. That's not problematic because burning oxygen-rich eliminates coking. When there is much more oxygen available then needed, all the fuel fully combusts. 

In contrast, burning hydrocarbons in a fuel-rich mixture is begging for coking. 

Staged combustion has a huge advantage because thrust-specific fuel consumption is the single biggest factor in keeping your mass ratio low. If you are throwing away fuel that doesn't contribute to thrust, you're throwing away specific impulse. 

[shynung]

@MathiLpHD

Note that I assumed identical propellant in the comparison above. If a methane FFSC turbine indeed runs cooler than RP-1^[dubious], then a methane GG system is going to run even cooler. That's because, again, a staged combustion cycle have the turbine exhaust into the main combustion chamber, which means the turbine preburner has to generate even more pressure to run the turbine. More pressure comes from more power (more propellant), and more power generate more heat.

Also, @sevenperforce, the GG system does trade off specific impulse for more thrust at a given cost.

Edited March 3, 2016 by shynung

[MathiLpHD]

44 minutes ago, sevenperforce said:

To answer why you can't simply keep increasing chamber pressure to get higher ISP: eventually, as you approach that upper theoretical limit, your fuel no longer has any more chemical energy available to increase the heat and pressure in the combustion chamber. 

Full-flow staged combustion has the disadvantage of requiring two separate preburners. However, this means the requirements on each one is lower, and since you already need separate pumps, this isn't much additional weight cost. 

And yes, the Russians did oxy-rich staged combustion with RP1. That's not problematic because burning oxygen-rich eliminates coking. When there is much more oxygen available then needed, all the fuel fully combusts. 

In contrast, burning hydrocarbons in a fuel-rich mixture is begging for coking. 

Staged combustion has a huge advantage because thrust-specific fuel consumption is the single biggest factor in keeping your mass ratio low. If you are throwing away fuel that doesn't contribute to thrust, you're throwing away specific impulse. 

But the coking temperature of methane is twice as high as the temperatures of a rocket engine, so that is not a problem with methane, right?

[sevenperforce]

9 minutes ago, MathiLpHD said:

But the coking temperature of methane is twice as high as the temperatures of a rocket engine, so that is not a problem with methane, right?

Not sure if that is the case for fuel-rich combustion or only for stoichiometric combustion. 

[MathiLpHD]

http://spaceflight101.com/spacex-launch-vehicle-concepts-designs/

"The advantage of the full-flow cycle is that the turbines operate at lower temperatures since more mass passes through them leading to increased reliability and a longer engine life which is particularly important to potential re-use of the engine. In addition, this engine design can deliver higher chamber pressures and improve the efficiency of the engine."

"Compared to RP-1, methane does not lead to coking of the engines which is a common problem with RP-1 that requires oxygen-rich combustion to limit coking, but creates a more corrosive environment."

[sevenperforce]

Yeah, did a little digging and it doesn't look like methane will coke even when run very fuel-rich.

If that's the case I'm not sure why they wouldn't ditch the oxy preburner and just run everything off the fuel-rich preburner. Hot oxy is bad stuff. 

[MathiLpHD]

Because they can run the turbines at lower temperatures and don't have to use exotic materials and get even higher isp from FFSC. And mixing both mixtures and light it causes the reactions to go faster, which might lead to higher thrust.

[K^2]

22 hours ago, MathiLpHD said:

4440/9.81 = 453s

So yes, it is possible to reach 400s with CH4.

Update2: For LH2/LOX it is ~575s.

The fact that LH2/LOX gives you 575s is precisely the sort of thing that should tell you that you'll fall short of 400s with CH4. But not by a whole lot.

[MathiLpHD]

The "bad" efficiency of LH2/LOX engines is mainly caused by the realy low density of H2. It's hard to run a turbine with H2 (e.g. on the expander cycle), so they can't get high pressure in the burning chamber. Methane has about four times the density of H2 and combined with LOX it is absolutly not hard to get high pump power and high chamber pressure. That's also the reason why they also use the oxygen-rich preburner. Because oxygen is much more dense so it can generate much more pressure through the turbines.

[shynung]

10 hours ago, MathiLpHD said:

Because they can run the turbines at lower temperatures and don't have to use exotic materials and get even higher isp from FFSC. And mixing both mixtures and light it causes the reactions to go faster, which might lead to higher thrust.

Except hot oxygen is terribly corrosive. The turbine exhaust pipes are the ones needing exotic materials.

If you really think staged combustion is the way to go, why not use the standard version rather than the full-flow version? If methane indeed runs cool enough to use on regular steel components, just use a bigger fuel-rich turbine to run everything together.

Edited March 4, 2016 by shynung

[MathiLpHD]

@shynung

It doesn't need "exotic" materials, it just needs corrosive resistant materials. I don't know if the standard zinc-steel of a car is good enough but i am not an expert for corrosion.

Because than you would have to use exotic materials for the turbine and still loose 10s of specific impulse because you need to power two pumps with one turbine so the pressure on the turbine is higher. And you can't just increase the size of the turbine without increasing the fuelflow because the turbine would loose efficiency and so the engine would loose efficiency. That's also the case if you throttle down an engine.

So i still think FFSC is the way to go for methane.

[Hannu2]

On 4.3.2016 at 8:46 PM, MathiLpHD said:

@shynung

It doesn't need "exotic" materials, it just needs corrosive resistant materials. I don't know if the standard zinc-steel of a car is good enough but i am not an expert for corrosion.

Because than you would have to use exotic materials for the turbine and still loose 10s of specific impulse because you need to power two pumps with one turbine so the pressure on the turbine is higher. And you can't just increase the size of the turbine without increasing the fuelflow because the turbine would loose efficiency and so the engine would loose efficiency. That's also the case if you throttle down an engine.

So i still think FFSC is the way to go for methane.

Standard steel from car is practically as resistant as cardboard if you have oxygen rich gas at temperatures typical in gas turbines. Burning through with oxygen jet is a very common method to cut steel in metal industry.

Corrosion resistance depends always on circumstances and chemicals. Anyone of so called corrosion resistant steels are not suitable for any purpose. There are hundreds (if not thousands) of different steel grades and those which are able to resist oxygen at more than 1200 K are very rare and expensive. Or in other word, exotic. Typically such dangerous conditions are tried to avoid in industrial processes and where avoiding is not possible, nickel alloys (even more exotic) are more common than steels as far as I know.

[RizzoTheRat]

The materials may be expensive but the technology is pretty mature. In standard gas turbines the turbine can be operating in an gas stream above the melting point of the material, and spinning at thousands of RPM generating massive stresses.  In some cases are grown from a single crystal to improve structural strength.

 

[shynung]

@RizzoTheRat The turbine, maybe, but the plumbing?