Oxygen Rich Engines?

[Titan Space Agency]

I am fascinated with the Nk-33 engine from the cold war. It has the second highest trust to weight of any flown. It also has a sea level ISP of 297 seconds. This is pretty high for a kerosene engine. The RD-170 family has an even higher ISP at around 309 seconds. These engines has something in common...

THEY USE OXYGEN RICH TECHNOLOGY!!!!!

Now i was wondering, is there any Hydrogen rocket engines that use Oxygen rich technology. It seems that it makes the engines preform with higher ISP.

[Ralathon]

I am fascinated with the Nk-33 engine from the cold war. It has the second highest trust to weight of any flown. It also has a sea level ISP of 297 seconds. This is pretty high for a kerosene engine. The RD-170 family has an even higher ISP at around 309 seconds. These engines has something in common...

THEY USE OXYGEN RICH TECHNOLOGY!!!!!

Now i was wondering, is there any Hydrogen rocket engines that use Oxygen rich technology. It seems that it makes the engines preform with higher ISP.

The reason oxygen rich fuel mixtures produce higher ISP is due to the lower mass per molecule. ISP is a measure of exhaust velocity and we produce that velocity by heating gasses. Lightweight gasses move faster at the same temperature. So using hydrogen gas at 2000K is more efficient than using Helium at 2000K. When burning kerosene and oxygen you're never going to get a perfect mixture, so you have some leftover oxygen and kerosene floating around in your hot exhaust. When you use an oxygen rich mixture you ensure that most of the leftover molecules are oxygen, which has a lower molecular mass than kerosene thus giving better Isp.

When you burn a LH2-LOX mixture the situation is reversed. Here hydrogen is the light molecule while oxygen is the heavy one. So if you use an oxygen rich mixture here you reduce the Isp. So you need to mix in some extra H2, and they indeed do this IRL.

[Guest]

Make that "a lot of extra H2". For optimum performance (SSMEs and the like), the LH2/LOX mass ratio is usually around 6. Mass ratio of a H2O particle is 0.125. (result of such combustion) Most of H2 in the rocket exhaust is actually just heated up and expelled from the nozzle. It burns anyway with atmospheric oxygen, that's probably why SSME (or any hydrolox engine) exhaust is so weird.

[Alchemist]

Actually, I once tried some calculations about the fuels and the standard LOX-kerosene scheme has slightly more oxygen than necessary for oxidation to CO + H2O.

So the oxygen-rich means the composition is closer to oxidizing the fuel to CO2 + H2O. What does it give?

Well, first of all that's more energy per mass of propellants (the key factor that limits theoretical ISP). However this extra energy is difficult to harvest.

For the same mass of fuel you get more energy but less particles. And more heat capacity. Result - higher temperature (slightly, the heat capacity eats part of the difference) and lower pressure at the same temperature and density (the actual engine might be rigged for higher work pressures). The question is getting the thermal energy out of this exhaust and converting it into kinetic - it's more job for the nozzle. Unless you make perfect nozzle for this system, you'll get worse ISP. That's why these were created much later than the simper solution.

Adding more H2 for LOX-LH2 has diametrically opposite effect - you get slightly less energy per mass, but it's easier to convert it to kinetic.

That's actually a sad story: Kuznetsov tried to pull this revolutionary design, but the reliability of NK-15 determined the failure of the entire N1 project, the upgraded NK-33 weren't even given a chance. Only years later Glushko perfected the scheme for his super-rocket (that flied only twice, but at least this time the engines weren't abandoned), and only recently the NK-33 were put out of the warehouse

[Temstar]

but the reliability of NK-15 determined the failure of the entire N1 project

Woah woah. That's in no way true. NK-15 is an extremely reliable engine, as evident by the fact that NK-33 built more than 40 years ago is still considered cutting edge hydrocarbon engine today and still in use. It's the N-1's 30 engine arrangement for the first stage that causes reliability problem. If you look at the reason why N1 failed:

First launch: metallic debris in the turbine of one of the engines, probably came from fuel tank

Second launch: a piece of slag in the fuel tanks or debris from a faulty fuel pump ingested by one of the engines

Third launch: fuel filter added to prevent the same type of issue as the first two launch, this time the problem was aerodynamic

Fourth launch: cause never identified

None of them were really caused by the NK-15. N1's layout was just a bit too novel and required test to iron out the bugs. It didn't help that N1 first stage was originally designed for 24 engines and they had to try to fit in six more since the spacecraft turned out to be overweight.

Glushko was Korolev's nemesis, so even though N1's fifth flight was to have the upgraded NK-33 and had a good chance of being a success, he was not going to tolerate Korolev having the last laugh on him with his magnum opus super rocket. And you have to remember that Glushko's thing was moon bases, so had N1 being a success under him he would have the tool to actually build his moon base. Even so he cancelled N1's fifth flight only to design his own super heavy lift vehicle 8 years later, after back tracking from his pro-hypergolic position.

Now as for oxygen rich, I don't think NK-15 / NK-33's exhaust is any more oxygen rich than other hydrocarbon engines. The "oxygen rich" part refers to it being a staged combustion cycle engine as opposed to the more common gas-generator cycle. Take for example Saturn V's F1 engines. To pump all that fuel and oxygen into the engine requires a very powerful turbopump. The turbopump is driven by what basically is a smaller rocket engine that burns some fuel and oxygen to generate hot gases, these hot gases spin up the turbine of the turbopump and provides the energy to pump vast amount of liquids to the main engine. Once the gas has going through the turbine blades it's just dumped overboard. Thus for a given level of fuel and oxidiser flow into a F1 engine, a small amount is spent driving the turbopump and thus do not contribute to thrust.

In a staged combustion engine, instead of dumping the turbopump exhaust out the side, the exhaust (which is still relative rich in oxygen, hence oxygen rich) is instead ducted so it flows into the firing chamber of the main engine, mixed with fresh oxygen and fuel and is fired again. Thus all of the oxygen and fuel that is feed to the engine eventually flows through the main firing chamber and out the nozzle. That's why staged combustion engine has higher Isp than gas-generator cycle engine - it's able to use all the fuel and oxidiser and shoot them out the back for thrust, instead of having to sacrifice a small amount to drive the turbopump.

The catch is oxygen is an oxidiser, which mean it really likes to react with things, weather that be RP-1 fuel or the metal walls of your engine. Oxygen as a hot gas (such as say, the turbopump's oxygen rich exhaust) is extremely corrosive and is likely to produce what rocket engineers like to call "engine rich exhaust". So in order to build a staged combustion engine you must have much better understand of metallurgy and gas dynamics. The Russians and only the Russians for some reason has mastered this technology. On the other hand Russians are pretty poor at building cryogenic engines unlike Americans (you can probably blame that on the stubborn Glushko and his insistence on hypergolic fuel), hence why no one has built a staged combustion cryogenic engine - the Americans don't have the staged combustion technology and the Russians don't know how to work with hydrogen.

Diagrams on staged combustion cycle vs gas generator cycle, notice where the orange turbopump exhaust goes:

Gas generator cycle

Staged combustion cycle

Edited March 28, 2014 by Temstar

[sgt_flyer]

Another feature (and difficulty) of staged combustion cycle, is the high chamber pressure. (70 bars for the f-1 engine, 145 bars for the nk-33. To support these kind of pressures and keep the weight of the engine down also require superior metallurgy. In the end, the higher chamber pressures gives a higher exhaust velocity for the gases - so better ISP. (Rd-180 has 266 bars of chamber pressure)

Edited March 28, 2014 by sgt_flyer

[Kerbal01]

Continuing this old thread, the RD-275 engine (first stage of Proton, uses 6 engines) does ORSC with Hydrazine, an extremely toxic/corrosive fuel. Way back in the mid 1960's. The Russians were crazy.

[fredinno]

3 hours ago, DarthVader said:

Continuing this old thread, the RD-275 engine (first stage of Proton, uses 6 engines) does ORSC with Hydrazine, an extremely toxic/corrosive fuel. Way back in the mid 1960's. The Russians were crazy.

So? The Titan used hypergolics too, and so did Ariane 1-4.

[Workable Goblin]

 

On March 28, 2014 at 5:48 AM, Temstar said:

The catch is oxygen is an oxidiser, which mean it really likes to react with things, weather that be RP-1 fuel or the metal walls of your engine. Oxygen as a hot gas (such as say, the turbopump's oxygen rich exhaust) is extremely corrosive and is likely to produce what rocket engineers like to call "engine rich exhaust". So in order to build a staged combustion engine you must have much better understand of metallurgy and gas dynamics. The Russians and only the Russians for some reason has mastered this technology. On the other hand Russians are pretty poor at building cryogenic engines unlike Americans (you can probably blame that on the stubborn Glushko and his insistence on hypergolic fuel), hence why no one has built a staged combustion cryogenic engine - the Americans don't have the staged combustion technology and the Russians don't know how to work with hydrogen.

That is absolutely not true. The RD-25, aka the SSME, is a staged combustion hydrolox engine, and a very, very good one (and also very, very expensive, as you might expect). The United States could do staged combustion in hydrogen just fine (despite the SSME's serious development difficulties, it worked very well during the Shuttle's lifetime); indeed, there were studies of kerosene staged combustion engines in the 1980s (like this one), and probably in the 1970s, too. It's just that there was no real incentive for developing more sophisticated lower-stage engines when existing designs could do the job at a lower cost. In the 1990s that changed a bit, particularly as EELV moved along, but it was still cheaper to just use Russian designs, so they did that on the Atlas V.

In theory the United States could build oxygen-rich staged combustion engines, it's just that no one has really felt like spending the billion or two dollars that would be necessary to set up production.

[fredinno]

9 minutes ago, Workable Goblin said:

 

That is absolutely not true. The RD-25, aka the SSME, is a staged combustion hydrolox engine, and a very, very good one (and also very, very expensive, as you might expect). The United States could do staged combustion in hydrogen just fine (despite the SSME's serious development difficulties, it worked very well during the Shuttle's lifetime); indeed, there were studies of kerosene staged combustion engines in the 1980s (like this one), and probably in the 1970s, too. It's just that there was no real incentive for developing more sophisticated lower-stage engines when existing designs could do the job at a lower cost. In the 1990s that changed a bit, particularly as EELV moved along, but it was still cheaper to just use Russian designs, so they did that on the Atlas V.

In theory the United States could build oxygen-rich staged combustion engines, it's just that no one has really felt like spending the billion or two dollars that would be necessary to set up production.

...Something LockMart is probably starting to regret.

[Temstar]

@Workable Goblin you are correct, my bad SSME is staged combustion. Is this the only one from the US though? I checked RS-68 and that's gas generator.

I'm guessing the SSME preburner exhaust is fuel-rich rather than oxygen rich though?

[GDJ]

If I recall correctly, the Soviets (Russians) had to develop a new class of high temperature Stainless Steel just to handle the oxygen rich fuel mix because the normal materials of the era just burnt away too quickly, which caused alot of failures (RD-180 comes to mind, IIRC).

I'm not sure of the metallurgical content of the Stainless Steel, but it was reputed to be a real PITA to machine properly.

Edited January 4, 2016 by GDJ

[blowfish]

I think there is a bit of confusion in this thread.  There is a distinction between fuel/oxidizer rich staged combustion cycles and fuel rich exhaust (you don't really ever have oxidizer-rich exhaust).  The two combustion cycle types refer to the gases running through the turbine, and not to the gas exiting the nozzle, which is almost always fuel rich.

The main reason to use oxidizer rich staged combustion as opposed to fuel rich staged combustion is that the fuel doesn't vaporize easily, as is the case with kerosene.  If your fuel is hydrogen, which does vaporize easily, you usually use fuel-rich staged combustion (though there are a few examples of ORSC engines burning hydrolox).

Edited January 4, 2016 by blowfish

[Bill Phil]

On March 23, 2014 at 11:42 AM, Ralathon said:

The reason oxygen rich fuel mixtures produce higher ISP is due to the lower mass per molecule. ISP is a measure of exhaust velocity and we produce that velocity by heating gasses. Lightweight gasses move faster at the same temperature. So using hydrogen gas at 2000K is more efficient than using Helium at 2000K. When burning kerosene and oxygen you're never going to get a perfect mixture, so you have some leftover oxygen and kerosene floating around in your hot exhaust. When you use an oxygen rich mixture you ensure that most of the leftover molecules are oxygen, which has a lower molecular mass than kerosene thus giving better Isp.

When you burn a LH2-LOX mixture the situation is reversed. Here hydrogen is the light molecule while oxygen is the heavy one. So if you use an oxygen rich mixture here you reduce the Isp. So you need to mix in some extra H2, and they indeed do this IRL.

Wouldn't this also change the thrust characteristics?

Another thing: Couldn't you mess around with some other attributes to lessen the impact?

[Kerbal01]

9 hours ago, fredinno said:

So? The Titan used hypergolics too, and so did Ariane 1-4.

Not ORSC.

[Workable Goblin]

10 hours ago, Temstar said:

@Workable Goblin you are correct, my bad SSME is staged combustion. Is this the only one from the US though? I checked RS-68 and that's gas generator.

I'm guessing the SSME preburner exhaust is fuel-rich rather than oxygen rich though?

Yes, there aren't any other staged-combustion hydrolox engines from the United States that I am aware of. There's a good reason for this, though; the SSMEs were required to be staged combustion to achieve the very high chamber pressures needed to have good performance from the ground to space, and this proved to be very expensive to develop and build. Other rockets, however, didn't (and don't) have the same boosted-sustainer design of the Shuttle, so they didn't need those chamber pressures. In particular, the RS-68 grew out of a program specifically designed to produce a cheaper booster version of the SSME for expendable applications, and so they went to a gas generator design to improve the characteristics (cost and TWR) that were important in that role. Similarly, staged combustion was unnecessary to achieve high performance in upper stage engines like the J-2X or RL-10/RL-60, so it again was not used in this role (in fact, there's a very clever trick that the RL-10 uses that gives you staged combustion-like qualities without actually having staged combustion, by taking advantage of some special properties of hydrogen. But it only works for small engines, not for engines with the SSME's thrust). In hydrogen engines, staged combustion has a very limited niche where it actually makes sense to use.

Also, you don't want to run hydrolox engines oxygen-rich. It produces an extremely high flame temperature and greatly reduces performance, as Ralathon explained. In the case of a kerosene engine, running oxygen-rich solves certain problems and can provide higher performance, but in the hydrogen case it just adds difficulties to the design.

[Shpaget]

On 3/23/2014 at 6:00 PM, Dragon01 said:

Make that "a lot of extra H2". For optimum performance (SSMEs and the like), the LH2/LOX mass ratio is usually around 6. Mass ratio of a H2O particle is 0.125. (result of such combustion) Most of H2 in the rocket exhaust is actually just heated up and expelled from the nozzle. It burns anyway with atmospheric oxygen, that's probably why SSME (or any hydrolox engine) exhaust is so weird.

I know this is an old post, but since the thread has already been resurrected...

The latests design of the external tank of the Space Shuttle carried roughly 630 t of LOX and 105 t of LH2, so it's the other way around. LH2/LOX ratio being burned is 0.167. There is 6 times more oxygen (mass) than hydrogen, instead of 8 time it would be required for stoichiometric mixture, so only 26 t out of 105 t of hydrogen is not burned.

The main reason for the rich mixture, I would assume, is to avoid having an oxidizing flame, which would be a problem for engines that are supposed to be reusable. Higher ISP is a side bonus.

Edit:
Oh, I forgot to say that, in this case, higher ISP doesn't necessarily mean higher dV. I haven't done the math, but if throwing out unreacted fuel was a good idea, they'd just open a valve at the bottom of a LH2 tank.

Edited January 4, 2016 by Shpaget

[blowfish]

18 minutes ago, Shpaget said:

I know this is an old post, but since the thread has already been resurrected...

The latests design of the external tank of the Space Shuttle carried roughly 630 t of LOX and 105 t of LH2, so it's the other way around. LH2/LOX ratio being burned is 0.167. There is 6 times more oxygen (mass) than hydrogen, instead of 8 time it would be required for stoichiometric mixture, so only 26 t out of 105 t of hydrogen is not burned.

The main reason for the rich mixture, I would assume, is to avoid having an oxidizing flame, which would be a problem for engines that are supposed to be reusable. Higher ISP is a side bonus.

Edit:
Oh, I forgot to say that, in this case, higher ISP doesn't necessarily mean higher dV. I haven't done the math, but if throwing out unreacted fuel was a good idea, they'd just open a valve at the bottom of a LH2 tank.

For the same fuel mass, higher I_sp does lead to higher delta-V.  For hydrolox, there is a definite performance benefit to running fuel-rich.  Obviously this only works up to a point, since the fuel needs to be heated - as you say, just venting LH2 won't create much thrust, but most real engines run at a lower mass ratio than stoichiometric (I've seen 4.5-6, as opposed to stoichiometric which is 8), because it increases I_sp.

[Shpaget]

There would not be the same mass. There would be 26 t less hydrogen, so if those engines ran at stoichiometric and managed to burn 100% the exhaust temperature would be higher than with rich mixture. Would the mass reduction and increased temp be enough to offset lack of H2 in exhaust, I don't know, but as I said earlier, are we sure the main reason to run rich is the ISP boost and not avoidance of oxidizing flame? I don't honestly know, I haven't done the math.

[blowfish]

2 minutes ago, Shpaget said:

There would not be the same mass. There would be 26 t less hydrogen, so if those engines ran at stoichiometric and managed to burn 100% the exhaust temperature would be higher than with rich mixture.

It's not useful to make a comparison like that, because decreasing fuel mass decreases delta-V even if you keep the mixture ratio constant.  It's more useful to think about keeping the propellant mass constant, but varying proportion LH2 vs LOX.  The only reason you might want to consider adjusting the overall mass is that running fuel rich reduces thrust somewhat.

8 minutes ago, Shpaget said:

but as I said earlier, are we sure the main reason to run rich is the ISP boost and not avoidance of oxidizing flame? I don't honestly know, I haven't done the math.

For hydrolox engines, yes, we are sure.  The mixture ratio is way below stoichiometric, and on some engines, the mixture ratio is actually varied to adjust thrust vs I_sp (example, the J-2 on the Saturn V's second stage adjusted its mixture ratio to increase I_sp once its full thrust output was no longer needed.  Whether there is another reason why engines with heavier fuels run slightly fuel-rich, I don't know.

[fredinno]

1 hour ago, blowfish said:

It's not useful to make a comparison like that, because decreasing fuel mass decreases delta-V even if you keep the mixture ratio constant.  It's more useful to think about keeping the propellant mass constant, but varying proportion LH2 vs LOX.  The only reason you might want to consider adjusting the overall mass is that running fuel rich reduces thrust somewhat.

For hydrolox engines, yes, we are sure.  The mixture ratio is way below stoichiometric, and on some engines, the mixture ratio is actually varied to adjust thrust vs I_sp (example, the J-2 on the Saturn V's second stage adjusted its mixture ratio to increase I_sp once its full thrust output was no longer needed.  Whether there is another reason why engines with heavier fuels run slightly fuel-rich, I don't know.

Kerlox engines is because LOX-rich burns hotter, and is more expensive to deal with.