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Convert Rocket Engines to Hydrogen or Methane


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Hello you all!


Some fabulous rocket engines exist already and use oxygen and kerosene in a staged combustion cycle (drawing), where all liquid oxygen is pumped to a huge pressure, burnt with little kerosene to achieve a reasonable temperature, passes through the turbine that moves the pumps, and burns the rest of the kerosene in a main chamber at high pressure. This most efficient cycle gives for instance the RD-170 its 8MN thrust (record) and 337s vacuum specific impulse (record with first-stage kerosene) and the RD-0124 its 359s specific impulse (record with kerosene). Better: the single slow turbo-pump makes cheap engines, and pressure makes them compact.





The hard development (pre-chamber and turbine in hot oxygen!) is already done; adaptation to hydrogen or methane seems much easier than a new design, with the method I propose. I had first described it there, and with more explanations and examples of uses later



but the present topic attempts to be clearer and more concise.


Essentially, I keep the oxygen pump, the pre-chamber, the turbine, the kerosene pump untouched or little modified and add a methane or hydrogen pump, in light blue on the schema:


In the new operation, the main kerosene flow to the chamber is replaced by the hydrogen or methane flow. The bit of kerosene is kept at the pre-chamber because it works, but methane could reach the huge pressure in one pump step as well; some kerosene flow can be kept to the main chamber, in tiny amount if it stabilizes the flame there, or in big amounts if one builds a tri-propellant engine (like oxygen and first kerosene, then hydrogen, which shall ease a single-stage-to-orbit).


Keeping some kerosene and its pump in addition to hydrogen or methane to may look complicated, but the parts already exist. A single stage hydrogen pump achieves the main chamber pressure, while the pre-chamber would require three stages with hydrogen.


Existing hydrogen engines with staged combustion have a fuel-rich pre-chamber. This needs all the bulky hydrogen to be pumped to huge pressure and produces very fast hot gas, both needing a fast turbopump with several stages - this is easier with the oxygen-rich pre-chamber. Such hydrogen engines had been tried at methane and tri-propellant operation, but kerosene and methane soot in a fuel-rich pre-chamber, ruining the efficiency: oxygen-rich is better for them.


Marc Schaefer, aka Enthalpy



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I'm not 100% sure I'm following everything you're saying but....


I'm not sure I follow your numbers. For starters, in one of your links you mention a combined fuel density *INCREASE*. Am I not reading something correctly or how do you propose replacing the majority of the kerosene with H2 but somehow increasing the density of the overall system?


Another question: You state that the hard part - handling a high temperature, oxygen rich environment - has been accomplished, but I see no mention of HOW it's accomplished. I mean, if the solution to nasty, expensive turbopumps is a pre-chamber made out of Rhenium, I submit that you're simply exchanging one expensive component for another. I'm not saying that's what you're proposing, mind you. I'm just pointing out that without any details, your claim comes across as smoke and mirrors.

Edited by InigoMontoya
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Methane and hydrogen are more bulky than kerosene, so the cycle achieves a main chamber pressure lower than the original. A new design would re-optimize all pressures and speeds but I want to keep most engine components, hence cycle adaptations are linked.


One possibility keeps all pressure ratios and temperatures, which maintains the gas speeds and keeps the turbine and its speed untouched. As much energy is harvested by the turbine per kg of heated oxygen gas; pumping liquid oxygen to the lower pressure takes less energy which is invested in the extra fuel volume. Gas density hence mass flows scale down as the pressure, reducing the engine's power. Methane keeps 222b chamber pressure from 245b at the RD-170 burning kerosene: in vacuum, the specific impulse jumps to 3526m/s=360s for 7.24MN :cool: thrust kept, while sea level operation provides 3209m/s=327s and 6.59MN, all with the original four small nozzles. This holds for the RD-180 (nearly a half RD-170, with two chambers and nozzles instead of four). The RD-191 (quarter RD-170 with one chamber and nozzle) lets more easily widen the nozzle exit.


Now if you look at the RD-170's turbopump, here carefully botched by myself:


it comprises one part with the gas turbine and the oxygen pump, and an other part with the kerosene pump. Each has a complete shaft with bearings and seals and a casing, so we can separate them at the dry interface to insert the additional pump in between. Even better: the oxygen pump, which is designed for the same cold and volume flow, would pump lighter methane to the main chamber pressure :cool:. In fact, pressure from the original impeller must be slightly reduced, both at oxygen and methane, for which I'd just reduce the impellers' diameter with a grinding machine :P. It also needs a bearing and seals at the methane pump section, new injectors at the chamber, and a methane booster pump - that's about all!


This is a small effort to gain 23s efficiency thanks to methane and enjoy a huge thrust.


Marc Schaefer, aka Enthalpy




The same parameters adaptation would work with hydrogen, keeping the turbine, but the low density would slash the pressure and thrust by a factor of 1.85, so sad. An other solution redesigns the turbine to extract more power from the gas expanding from the pre-chamber at original pressure to the main chamber at a lower pressure; more power allows a main chamber pressure reduced by 1.44 only, and the full pre-chamber operation keeps the full oxygen mass flow.


Here's an example, this time with the smaller RD-0124 (Soyuz' improved third stage):


again, the oxygen and kerosene sections have a clean dry interface where the additional pump section is inserted. The parts are approximately to scale: light hydrogen needs a larger impeller on this slow shaft, but its peripheral 540m/s are usual and its 0.26m diameter still reasonable. The hydrogen inlet and feeder are big to avoid a separate booster pump.


The original four nozzles provide a big expansion for vacuum operation. With hydrogen at 109b, the adapted engine would boast 4382m/s=447s and 306kN:cool:: efficiency from the best hydrogen engines group, with more thrust, through a limited adaptation effort.


Marc Schaefer, aka Enthalpy




I go to bed, but will come back. With answers for Iñigo (¡Gracias por tu interés!), more engine descriptions, and examples of use at new launchers or for retrofit, including as side boosters for Ariane 5.

Edited by Enthalpy
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OK, so you gain 23 s of Isp. But what does it do to density Isp? More to the point: Since you're using a significantly less dense fuel, does the increase in size in fuel tankage, related structure, and new-found fuel tank cryogenic requiremenst make it worth it? Mind you, for 2nd stage and above flight, I'm sure it does but then, I'm not aware of anybody burning kerosene for 2nd stages. I presume you're talking 1st stages. For 1st stages, density Isp seems to be a better metric for performance than simple Isp.

Edited by InigoMontoya
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Hi Iñigo and everyone!


In the other thread, "Higher combined propellants density" only compares with straight hydrogen+oxygen. With this reference, the few % kerosene gain few % density, and the added tank can fit in unused room. But if referring to a dense propellant like kerosene, density is clearly less good, and is very near to normal hydrogen+oxygen or methane+oxygen, sure.


Isp is much more important than density for space rockets. Isp*density is used for missiles, where mass ratio is less important and volume is limited - and even then, this criterion is voluntarily biased by solid fuel suppliers to favour their junk. Isp*density is the N*s a fuel provides from a given tank volume, as if fuel mass were unimportant!


Kerosene is used frequently on 2nd and 3rd stages, for instance by Soyuz, Zenit, Falcon-1 and Falcon-9... It offers an excellent blend of performance even for upper stages, but hydrogen is so seducing that it tempts everyone for upper stages, including Soyuz and Falcon. Hydrogen at a first stage is debated as it makes the launcher bulky, but tank mass isn't necessarily an obstacle - chosen for Delta IV. And imagine a stage like Proton: the smaller central oxygen tank could transmit the thrust to the upper stages, while lateral hydrogen tanks would only support themselves and be light.


The sketched tanks are to scale. I'll put here as well some launcher designs.




Hot oxygen is a harsh environment for a turbine - not so much for the pre-chamber, which is actively cooled by liquid oxygen, including the injectors at the RD-170. This is what Google translates from Russian at Lpre.de about the RD-170:

To eliminate the fire and destruction of parts of gas turbines tract, the design incorporates nickel alloys, including superalloys for the hot gas lines. The stator and the turbine exhaust tract forcibly cooled with cold oxygen. In places small radial clearance or end-uses various heat-resistant coating (nickel for rotor blades and stator, the rotor cermet), as well as silver or bronze items, excluding the fire even with a possible touch of rotating and stationary parts turbopump unit.


So: the usual nickel superalloys, with a nickel or a ceramic coating. It certainly needed much experimental development, the reason why I like reusing this existing technology, but material cost looks reasonable - consistently with the RD-180's low cost: I've read figures like 10M$ for the original 4MN kerosene engine, as compared with 20-30M$ for a 3MN RS-68. A limited adaptation getting 8MN from an RD-170 with methane or hydrogen Isp is hence attractive.


Marc Schaefer, aka Enthalpy




A single impeller can pump hydrogen to 100 or 150b at once with common materials. The slow shaft of an oxygen-kerosene turbopump implies a big impeller, as on the sketch for the RD-0124. Maybe a hydrogen impeller adapted from the RS-68 would fit the smaller RD-191, but not the RD-170, which would have needed D=0.83m. Several stages would reduce the diameter but are just too sad for only 150b.


I believe my Pelton-Schaefer pump :rolleyes :is a solution. Mentioned and later detailed there:

http://saposjoint.ne...6&t=2272#p27829 on Fri Jul 16, 2010

http://saposjoint.ne...start=40#p33837 on Tue Aug 30, 2011


Just as a Pelton turbine, its blades run at V, and the liquid at nearly 0 and 2*V:

http://en.wikipedia....ki/Pelton_wheel - other images at De and It wiki, the source of this one:


which is the trick, because 150b with hydrogen need 2*V=650m/s and V=325m/s is easy for the wheel.


As a Pelton turbine as well, the wheel is surrounded with gas or vapour, and a thin liquid film flows on the blades, fast but at low pressure. No bubbles need to collapse, and Pelton turbines last for years at 150b dynamic pressure.


As opposed to a turbine, the pump takes liquid at small speed and pressure and accelerates it to 2*V, which is the sought overhead, converted into pressure by the discharge channel (the cut is through the blades near the centre of this sketch but equatorial elsewhere):


the blades scoop liquid much like a spoon scoops ice cream or a chipping tool cuts material from a part. This allows also a big volume flow in a compact pump, more so with a booster pump. Like: D=0.38m for an RD-170, with 40mm blade height - easy fit on the original turbopump.


This pump brings no throughput stability, that is, its volume flow is independent from the discharge pressure until goes mad. A fast active control must regulate the flow at the inlet.


Marc Schaefer, aka Enthalpy




A strong and efficient first stage opens new possibilities, for instance split the rocket performance to geostationary orbit (13600m/s) evenly between just two stages, with comfortable inert mass. (Click on the images to display at true size)


As compared with the somewhat more capable Ariane 5, mass saved by the more efficient propellants lets bulky hydrogen look reasonable. The four configurations bring the same performance.

Boring figures starting there http://saposjoint.ne...t=2272&start=60 if you like them.


Now if a retrofit shall improve the performance of an existing launcher, sure it gets bulky:


Unusual and looking odd: the hydrogen tanks are at the sides of the oxygen tanks. All are pushed by the engines, only the oxygen tanks transmit thrust to the central stage and the upper composite. I took 100kg tank mass per ton of propellants, but it's very pessimistic then.

The left design keeps lift-off mass for compatibility, while the centre one makes good use of the two RD-170: 20t more to Leo than if burning kerosene.


More payload using at the second stage an RD-191 modified for hydrogen:


still compact despite hydrogen everywhere.


Even more payload with an RD-170 at the second stage, modified for hydrogen:


but this time the first stage burns methane and oxygen, pressure-fed from steel tanks, to save money.

These are the sailback boosters I describe there:


which shall be reused after splashdown in the Ocean.


Marc Schaefer, aka Enthalpy

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I have read most of the text here in the thread, but I cannot figure out the reason for the extra thrust. Is it the other fuel (hydrogen and/or methane vs. kerosene) or the oxygen rich (instead of fuel rich) pre-combustion?


Hydrogen and oxygen are nothing new. The Space Shuttle uses used it too. Kerosene is just cheaper and easier to handle, and that's why it's used in many rockets.

Regarding the oxygen rich chamber, I cannot figure out any thermodynamic reason why this would be more efficient.


Can you give a summary of the point you try to make, instead of such a long story? It's a little difficult to follow.

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Hi CaptainPanic and all, thanks for your interest!


The extra thrust was unexpected indeed. It doesn't come from more ejected mass, as the oxygen throughput is kept in the best case, and O2:RG1:H2 mass ratio is 1000:28:167 instead of the initial O2:RG1 mass ratio of 1000:427. It comes from the greater ejection speed, like 4023m/s instead of 3305m/s for the RD-170 keeping the nozzle exit area.


Hydrogen engines are much more efficient but more difficult, and the strongest existing today (RS-68) pushes 3MN, though it is meant for first stages (Atlas V). The possibility to convert an existing kerosene engine enables new performance: for instance, an Ariane 5 with two RD-170 boosters at maximum lift-off mass would put 50t into low Earth orbit, but 70t with hydrogen. An other example goes to geostationary orbit in two stages, which would be very bad with kerosene. As well, all highest-exhaust-speed hydrogen engines today use an expansion cycle that lets them push 100kN (RL-10, RD-0146...) and soon 180kN (Vinci); staged combustion as opposed doesn't limit thrust and the RD-0124 for instance would push 300kN with a similar exhaust speed.


Staged combustion provides better performance among the powerful engines because all of one propellant passes through the turbine, not just a derived fraction of the propellants, and this achieves a higher chamber pressure. Oxygen-rich is the only way to burn properly a hydrocarbon prior to the turbine, as this turbine limits the temperature and impose very detuned mixtures; very fuel-rich would let soot even methane, and this soot is lost for combustion. Fuel-rich is usable and used with hydrogen. Adapting hydrogen engines to methane produced disappointing results, I believe for this very reason.


Now, if the goal is a hydrogen-oxygen engine:

- A fuel-rich pre-chamber achieves about the same performance. This is by chance, as many parameters differ: the amount of propellant passing the turbine, the volume of propellants pumped to the huge pre-chamber pressure, the worse efficiency of the hydrogen pumps and turbines.

- Oxygen-rich brings fire risk at the materials, but hydrogen-rich needs huge turbine and pump speeds, despite using several stages.

- Kerosene engines exist already and are cheap, even for 8MN, hence my interest in converting them to more efficient fuels through limited modifications.

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