The Russian Raptors: the RD-0162/0164 methalox engine family

[Hyperion5]

Amidst the discussion of Spacex's announced new Raptor engine, a few of you may have noticed that the Raptor is constantly being compared to one under-design Russian engine family: the RD-0162/0164 engine family.  This engine family began life in 2002 as part of a European-Russian program named "Volga": http://www.kbkha.ru/?p=8&cat=11&prod=59.  They have only been under design as part of the Russian MRKS program since 2006.  Like the Raptor, these engines are meant to be highly reusable (up to 25 uses per), will burn liquid natural gas (AKA methane) with liquid oxygen (methalox), and they feature a medium-pressure staged combustion cycle, and they're even similar in thrust.  The RD-0164 in particular is rated at 300 tonnes-force in a vacuum, giving it the same vacuum thrust as a Raptor engine.  Elon Musk was even asked, due in no small part to KBKhA's methalox engine expertise, whether he intended to work with the Russians (http://shitelonsays.com/transcript/elon-musk-lecture-at-the-royal-aeronautical-society-2012-11-16 --use control+f to find the 2nd use of "Russians"). 

"Are you going to work with the Russians?" 
Elon Musk: "No.  We might hire a few Russians but..yeah." 

Ironically the RD-0162 will probably be beaten to the test stand by the Raptor engine, despite the Russians' lead time.  Since 2012 they've started work, probably at a low-level, on a smaller, lower-pressure derivative called the RD-0162SD.  They have already test-fired a methalox version of the RD-0110 called the RD-0110MD.  The RD-0162/0164 engine family is so promising that TsSKB Progress, makers of the Soyuz rocket family, have begun imagining an entirely new Soyuz family powered by the larger RD-0164 variant called the Soyuz-5 family: http://www.russianspaceweb.com/soyuz5.html.  Even more entertainingly, TsSKB has proposed a super-heavy lift rocket family using large numbers of these engines: http://www.russianspaceweb.com/stk.html.  It's a rather interesting coincidence, given Spacex will likely be using large numbers of Raptor engines for an HLV family of their own. 

The RD-0162, while not matching SC kerolox engines in impulse density, gets close to matching an NK-33 engine in thrust:weight ratio when maxing out its thrust and features excellent Isp.  These engines can push their thrust to 133% of nominal, allowing their thrust:weight ratios to jump to ~130. 


RD-0162 basic parameters   
Thrust in tonnes-force (kN)   203.9 (2000)   
Specific impulse, kgf · s / kg (m / c)
Sea level: 321 (3149)
Vacuum:    356 (3492)   
Chamber pressure, kgf / cm ² (MPa)   175 (17.1)   
Uses:   25   
Max Thrust Rating in %:   133   

Propellants: lox + LNG

Weight in kg:   2100   
Motor size , mm
height: 3550
diameter: 1650   

Recently, I challenged one of our forum members, Malu, to test his engine calculator against the known specifications of the RD-0162 and see what specifications his calculator suggested for the engine in contrast.  The results were interesting to say the least. 

This was one tough challenge and I'm not done with a solution, however here's a quick update after trying some alternative ways to model the RD-0162.

With my standard approach of using braeunig's graphs for Methane/LOX rocket engines, I came up with 1.94 m diameter nozzle using 2.9/2.93 mixture ratio 358s isp Vac (100% throttle) and 324s Isp SL (133% throttle) (calc number 2), this is the answer to your question.

However, intrigued by this I continued to unravel this mystery. In calculation number 3, I assume variable throat area, using this it's easy to find a solution where restrictions on nozzle diameter, thrust and sl/vac ISP is met. Googling seem to indicate there might be ways of achieving "variable throat area", however, it seems extremely difficult.

In calculation number 4 i keep also the throat area constant (more realistic), however, this puts a limitation on the 100% thrust level @ 322 klbf (133% thrust fullfil the requirement of 449 klbs, so this might be plausible if all citations of "max/SL thrust are meaning the 133% thrust level) - which makes sence since this level likely will be used at lift-off and thus is of interest when designing the rocket.

However, then I found this russian paper on the RD-0162 (РДО162) and using the translation feature in preview I could understand some of it. I also dug up various other references to specs and they all point to a 3.5:1 mixture ratio. In all my calculations, starting off from braeunig's graphs I have used ~2.9. Guesstimating what it should be for mixture ratio 3.5:1 I can model an engine which match all specs at 100, see calucation number 1.

The specs my model, as is, would predict the following specs for RD-0162;

Note; The correct chamber pressure and adiabatic flame temperature using mixture ratio 3.5:1 specified pressure, are not known. Guesstimates are used.

	РДО162 @ 100%	РДО162 @ 133%
Nozzle Diameter, m	1.65	1.65
Throat area, m^2	0.069217	0.069217
Chamber Pressure, Mpa	17.1	23.4
Chamber Temperature, K	3940*	4200*
Thrust, SL, lbf	449'000	629'000
Thrust, Vac, lbf	497'000	677'000
Isp, SL, s	321.9	338.9
Isp, Vac, s	356.2	364.7

*) Chamber temperature, molar mass and specific heat are not known to me at mixture ratio 3.5:1 and given pressure, I'm looking into as next step using Cantera to understand what's going on when using this mixture ratio, since braeunig's graphs only goes to 2.9:1.

In all examples given, the engine is very over expanded at 100% thrust @ 1atm ambient pressure (sea-level). In calculation 2 and 4 it seems to be below the limit of what's acceptable and must run at higher thrust than 100% at sea level. In almost all calculations, the 133% throttled engined with it's higher engine pressure, shows better SL and Vac ISP, even though it's under-expanded at SL. I believe the point of this engine is to operate mostly at 100% for reusability with reasonable Vac performance, while the 133% thrust-level is for initial lift-off to minimise gravity losses while keeping engine wear and tear as low as possible.

It's also an interesting design consideration to make engine nozzles smaller than SpaceX currently do; trade thrust penalty from over expansion at sea level against gravity losses at higher altitudes (smaller nozzle => require throttle down w. higher gravity losses but better Isp @ Vac/low pressure).

Feel free to comment on anything regarding these engines or their relation to the rockets they've been proposed to power. 

[MP99]

However, intrigued by this I continued to unravel this mystery. In calculation number 3, I assume variable throat area, using this it's easy to find a solution where restrictions on nozzle diameter, thrust and sl/vac ISP is met. Googling seem to indicate there might be ways of achieving "variable throat area", however, it seems extremely difficult. 

I've been wondering about this.

Could the throat use Shape Memory Alloy? http://en.m.wikipedia.org/wiki/Shape-memory_alloy - see the one way effect.

It would start in stretched (low temperature) form, IE a wide throat. Cryogenic propellant would flow through cooling ducts to keep the alloy below it's transition temperature in the lower atmosphere.

Once it is safe to increase the area ratio without danger of under expanding the nozzle, the flow of coolant is reduced to allow the alloy to go above it's transition temperature. This will cause it to revert to it's pre-stretched shape, IE a smaller throat diameter, increasing the area ratio, and providing a step change in the Isp.

I assume that test firings of the engine would maintain the SMA below it's transition temperature, so it will deform to it's vacuum form for the first time in flight.

I don't know how the Russians plan to recover their stage. But for SpaceX's recovery method, the post-MECO Boostback & Reentry burns will be in vacuum, so this is not a problem. However the landing burn will be at SL, which would cause the nozzle to be under expanded without a boost to the chamber pressure. Perhaps this is acceptable for the brief burn involved?

For the engine to be reused, the throat would need to be stretched back to it's Sea Level diameter as part of preparations for the next flight. If there is a hot fire test before the next flight, that could verify the accuracy of the re-sizing of the nozzle, as part of the startup checks. Of course, these are also part of the hold before release system on launch, but the hot fire allows humans to go over the results in more detail to sign off the re-configuration.

Thoughts?

Cheers, Martin

[Robotbeat]

Changing throat diameter for a circular (not annular) throat is nearly impossible and certainly isn't feasible. There ARE ways to change throat area for other geometries, though I doubt it's worth it.

Also, they should run the numbers for a 3m and 6m diameter nozzle and vacuum conditions. I could run them in RPA myself, I guess, but it'd be nice to compare. For an upper stage engine, 3m or even 4m may be appropriate, though it ends up getting pretty long... Which can be solved with an extendible bell, like on the RL-10b-2

[R7]

OK thank you forum software for necrothreading warning but I do believe hrissan's great find deserves to be stored in the right thread;

No mention of other details.

Exactly, so wondering where the dual-turbo comment came from.

http://www.kbkha.ru/?p=4&cat=51#souz

Двигатель содержит низкотемпературный двухконтурный газотурбинный тракт с двумя турбинами (окислительный газогенераторный контур и восстановительный безгенераторный контур).

It is dual turbine.

What is interesting: 2 turbines, the bigger one is driven by gas-generator,  the smaller one by evaporating methane in cooling contour.

It is not clear from description which turbine drives which pumps, but after reading this work

http://www.dissercat.com/content/povyshenie-effektivnosti-sistemy-podachi-topliva-kislorodno-metanovogo-zhrd-s-dozhiganiem-vo

It becomes apparent that bigger turbine drives both main pumps, and the smaller one drives inlet "booster" pumps.

So not a FFSC engine?

This engine is gas-gas. Full flow of methane must be evaporated in thick channels around main chamber (and flimsy tube you see around nozzle).

So if you have all your methane vaporized already without additional gas generator, and you have even excess energy in it to run turbine for inlet pumps, why not? So you have an elegant solution, the engine may be very similar to kerosene engine family.

Single main turbine running on oxygen-rich gas generator has both main pumps on single shaft, does it sound familiar? This also sounds lightweight.

Very interesting cycle indeed. While relying on google translation I got the idea that expander side drives a booster pumping stage similar to mechanically driven booster pump in NK-33. It raised the pressure for the preburner part of fuel flow.

A kingdom for proper pneumohydraulic scheme!

[TomH]

H5:

Two questions:

What effect do you see the ruble crash having on this research?
Do you believe they will use unrefined LNG rather than pure CH₄?

[russianhalo117]

H5:

Two questions:

What effect do you see the ruble crash having on this research?
Do you believe they will use unrefined LNG rather than pure CH₄?

Unrefined LNG is not an efficient option.