Staged Combustion Rocket Engines

[Enthalpy]

Hello everybody!

 

Staged combustion lets one propellant pass fully through a pre-chamber (or gas generator) at high pressure, adds a bit of the other propellant to burn it partially to a moderate temperature, and pass the big flow through the turbine. This gives more power to the pumps, and the higher pressure in the main chamber makes a more efficient engine.

I've much simplified my obscure sketches here under; the reader may imagine in his clear mind:

• Each propellant passes first through a booster turbopump.
• A second impeller brings only the small fraction of auxiliary propellant to the pre-chamber pressure. This saves power, the gain is notable.
• A propellant flows through the cooling jacket. Preferably at nearly the chamber pressure.

Staged combustion has been used on many hydrogen engines. As opposed, hydrocarbon fuels (including methane) would soot with so little oxygen, and oxygen consitutes most their propellant moles anyway, so the pre-chamber is oxygen-rich for hydrocarbons; the difficulty of hot oxygen has been addressed only by the Soviets and Russians up to now, resulting in the superior Rd-180 engine for instance.

 

 

As an exception, methylamine CH₅N would not soot with little oxygen. One might also try ethylene diamine dissolving much guanidine and maybe some methylamine; there are very few possibilities.

Because its hot gas is lighter, and nitrogen adds moles, methylamine achieves almost the same main chamber pressure as oxygen with Rg-1, and this better fuel gain 5s specific impulse, with a fuel-rich pre-chamber. The comparison conditions are imperfect but fair: 700bar and 600°C=873K in the pre-chamber, 74%-70%-88% efficiency at the turbines-pumps-injectors, 1.25 mechanical power margin. Though, I doubt about Codata's Hf298Liq = -47.3kJ/mol.

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The next step is a full-flow staged combustion. It was considered for hydrogen-oxygen, and I believe never built; methylamine enables it. Two pre-chambers burn separately all the oxygen and fuel at moderate temperature: more gas in the turbines raise the main chamber pressure to further gain 4s specific impulse. It takes two turbopumps, but the sealing joints are easier.

 

 

The fuel turbine has some power left and can optionally help rotate the oxygen pump if coupling the shafts:

• This raises the main chamber pressure from 312 to 334 bar.
• It gains only 1s specific impulse.
• Integration and seal joints are more difficult.

5s or 9s improvements do bring an advantage, but I wouldn't go to methylamine:

• It's very volatile (bp -6°C), flammable, toxic and caustic as an amine.
• It decomposes irregularly at heat or banal catalysts, without oxygen, producing hot gas.
• This implies oxygen in the cooling jacket. It has already been done but not gladly.
• Oxygen-rich staged combustion was developed five decades ago. It must be feasible now with better alloys, ceramic coatings, despite computers. Fuel-rich methylamine isn't required for staged combustion.
• A molybdenum turbine brings about the same improvement. Light big nozzles of SiC or niobium, also.
• Replacing Rg-1 by methyl- or cyclopropyl- substituted azetidine, diazetidine, diaza-spiroheptane (choose for liquid range), as well.
• Oxygen in the cooling jacket enables other, more efficient fuels.
• Maybe a combination of nonvolatile amines can replace methylamine?

Marc Schaefer, aka Enthalpy

[Enthalpy]

In the pre-chamber of a staged combustion engine, the propellant ratio is extremely de-tuned so the gas temperature fits the turbine. This makes the flame difficult to start, keep and re-start.

The necessary oxygen amount at once would quench the flame, so instead, a small flow of oxygen first burns the fuel in a balanced, hot, slower flame; then the big flow is injected to obtain much lukewarm gas - there the combustion stops.

Cyclopropane and methane (hi there), and all dense fuels, have the same difficulty as kerosene; only hydrogen is easier. This applies also to a fuel-rich pre-chamber where possible - so the sketch can generalize to "minor propellant" and "major propellant".

The Rd-170 engine family brings the kerosene and both oxygen flows at each injector individually; due to the dilution and the gas speed after the final injection, the flame can't spread from an injector to an other. As Energomash states, "each injector is a combustion chamber". Some trimethylgallium and trimethylaluminium, hypergolic with oxygen, flow before kerosene to light it. This design has advantages like easily cooled parts, but trimethylgallium is dangerous and the engine would be difficult to restart in flight.

I suggest instead to let the hot flames communicate between the injectors, upstream of the main oxygen flow.

One or few igniters then suffice, avoiding the pyrophoric propellant. One example is a Diesel glow plug, optionally adapted to rocket engines as I describe there
http://forum.nasaspaceflight.com/index.php?topic=27308.0 on 01/13/2012
it likes oxygen-rich flames, has a significant resistance to heat, and is insensitive to soot.

The sketched whirl injectors aren't necessarily the best choice. Parts of the pre-chamber must be cooled like a main chamber. And because the small flow flame has more room available, the pre-chamber might be downsized.

Marc Schaefer, aka Enthalpy

[Enthalpy]

In the full-flow staged combustion of oxygen and methylamine (or hopefully a combination of less volatile amines), no oxygen needs to be injected in the amine pre-chamber. The recomposition of the amine produces enough heat to achieve the 312 bar permitted by the oxygen side. This needs a stable reaction of the amine in the pre-chamber; catalytic recomposition of methylamine is known to be unstable, but maybe the hot pre-chamber solves that.

[Enthalpy]

If seeking the same moderate pressure and performance, a hydrogen staged combustion cycle is simpler than a gas generator cycle.

 

 

A single stage hydrogen pump after the (not represented) 20b booster pump achieves 123b in the pre-chamber, and the smaller pumping power leaves 103b in the main chamber, which gives the same performance as a gas generator cycle.

The hot gas' maximum expansion speed can be shared as 691m/s and 421m/s in the single-stage turbines that power pumps with 528m/s and 141m/s tip speed.

Single stages simplify turbines and pumps. Gas generator cycles exploit much faster hot gas through several stages.
We can also accept some liquid leaking into the hot gas if this flows in the drawn direction, which makes seals easier.

 

Matc Schaefer, aka Enthalpy

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Few days ago, Ariane's member states agreed on a new definitive design for Ariane 6, with one solid and two hydrogen stages. It's probably the best architecture without designing a new liquid engine.

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The Vulcain 2 kept from Ariane 5 is an old and not quite simple design. While better industrial organization and production methods can make it cheaper than now, at some point it will reach a barrier. The staged combustion design suggested just above may be cheaper than the gas generator cycle. It differs from the MC2000E option for Vulcain 3 by its lower pressures and higher pre-chamber temperature.

An other option would push the central stage with 7 Vinci chambers with D=1.3m nozzles and one common set of turbopumps and actuators. The chambers are kept as they are, the turbines and pumps are sqrt(7) times bigger, the angular speeds sqrt(7) times smaller like the resonant frequencies. This gains 10s isp over a Vulcain 2, saves the gas generator, and Vinci's turbopumps are much cheaper.

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I hope the designers of the integration and launch buildings will leave room for a wider body, and at the launch pad for more solid boosters, to keep the future open.

The sunheat engine brings more to satellites and space probes than any variation of chemical rocket engines
http://www.scienceforums.net/topic/76627-solar-thermal-rocket/

and it prefers wide fairings, so extra room at the integration and launch buildings would be smart.

Marc Schaefer, aka Enthalpy