Exotic Pumping Cycles for Rocket Engines

[Enthalpy]

Hello you all!

Many rocket engines pump liquid propellants to their combustion chamber, and varied cycles are used to power the pumps, the best known being:
Gas generator http://en.wikipedia.org/wiki/Gas-generator_cycle_(rocket)
Staged combustion http://en.wikipedia.org/wiki/Staged_combustion_cycle_(rocket)
Expander http://en.wikipedia.org/wiki/Expander_cycle_(rocket)
but here I'd like to describe uncommon cycles, which may be new and of my invention.

 

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The first sketched cycle realizes the cracking (hydrogenolysis) in a pre-chamber of a hydrocarbon with hydrogen, which produces methane, some excess hydrogen, and enough heat that the following turbine powers all pumps that achieve for instance 440b in the pre-chamber and some 200b in the combustion chamber. An amine, imine, nitrile can replace the hydrocarbon:

 

 

Logically, fuel density and specific impulse are between methane-oxygen and hydrogen-oxygen, thus filling the gap between kerosene and hydrogen engines:

 

 

The pre-chamber and turbine run with fuel-rich hot gas, far easier than the staged combustion of hydrocarbon fuels which was oxygen-rich to avoid soot.

Marc Schaefer, aka Enthalpy

 

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Now, this other cycle decomposes an endothermic oxidizer in a pre-chamber to obtain hot gas for the turbine. All the oxidizer flow decomposes and continues to the main chamber.

 

 

It shall be simple and reliable:

• No mix is required at the pre-chamber, which runs at a safe fixed temperature (but in oxidizing gas, true)
• Pump and turbine speed is moderate
• If you have some limited pressure in the tanks, opening two valves starts the engine, and the attitude control can be pressure-fed
• The oxidizing gas lights the fuel by its mere temperature
• The moderate decomposition temperature permits a sort of glow-plug igniter

 

This cycle burns storable propellants and is more efficient than tetroxide and toxic hydrazine, like 350s with a good expansion, and its tanks are lighter.

One oxidizer is Mon-33, or 33% NO dissolved in 67% N₂O₄. It freezes at -107°C, so if paired for instance with 2,4,6-trimethyl-tridecane (freezes at -102°C), they stay indefinitely on Mars or an asteroid or the Moon just in white tanks. Less NO lowers the vapour pressure at terrestrial temperatures, lowers the pre-chamber temperature, and loses little performance.

Hydrogen peroxide H₂O₂ is used that way by the RD-161P engine:
http://esamultimedia.esa.int/docs/EMO/LPE.pdf
but to this option, I prefer a liquid oxygen staged combustion, because peroxide isn't really storable and is dangerous.
http://www.gkllc.com/lit/gk-authored/AIAA-2004-4146_Field_Handling_of_hydrogen_peroxide.pdf

Marc Schaefer, aka Enthalpy

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And that other cycle recomposes in a pre-chamber a mix of amines to produce methane, nitrogen, little hydrogen, and heat:

 

 

Few amine mixes don't soot. I consider Ethylenediamine dissolving 358:1000 of Guanidine. The pre-chamber, turbine and pumps work then at comfortable temperature and speed, and the hot gas is fuel-rich.

• Ethylenediamine NCCN or C₂H₈N₂ is liquid between +9°C and +116°C, nearly safe, and its recomposition produces a temperature over the decomposition and the atmospheric autoignition, helping to recompose in the pre-chamber. But it needs some hydrogen supplement.
• Guanidine N=C(N)N or CH₅N₃ melts at +50°C and decomposes at +150°C. If 358:1000 dissolve in Ethylenediamine, sooting is overcome, and they hopefully lower the freezing point. Note: I've taken published -56kJ/mol heat of formation but doubt it; +56kJ/mol would soot.
• Methylamine NC or CH₅N boils at -6°C and its recomposition is reputedly unstable, but 240:1000 in Ethylenediamine overcome sooting and should lower the freezing point. Guanidine and Methylamine dissolved both in Ethylenediamine may combine the best freezing point and vapour pressure.

The sketch suggests an optional pressure-fed attitude control, but since only oxygen gives good performance here and is more difficult to store, this cycle would rather fit some launcher's lower stages. Attitude control uses to gimbal engines there, and booster pumps accepting a low input pressure save tank mass. This cycle is as good as oxygen-kerosene in a staged combustion but far simpler.

Marc Schaefer, aka Enthalpy

 

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For the cycle that recomposes an amine in a pre-chamber, Methylamine (MA) is corrosive and boils at -6°C, yuk; to avoid soot, Ethylenediamine (EDA) would need 100:23 of dissolved Methylamine (yuk) or 100:36 Guanidine, but the solubility and the resulting viscosity are doubtful.

Alternately, Ethylenediamine could dissolve just 100:3.2 of Ammonia to suppress sooting. I've taken the heat of formation of liquid ammonia for want of the solute. Ammonia is nasty, but 3% should produce little vapour pressure - less than 100:23 Methylamine.

 

 

Once Ethylenediamine contains the Ammonia bit, one can add Guanidine, Aminoguanidine or Diaminoguanidine.
http://en.wikipedia.org/wiki/Aminoguanidine (beware 1720kg/m³ matches software divination)
Aminoguanidine=Pimagedine was tested at 300mg/day as a drug, so it isn't very toxic. More aminated Guanidines bring heat (smoother decomposition; molybdenum turbine if needed) and performance but can't replace Ammonia - at the non-soot minimum amount in the table.

In the table, added Guanidines only keep the exhaust speed of Ethylenediamine + Ammonia at best, because this comparison is at identical pressures: 700bar in the recomposition pre-chamber, 300bar for combustion in Oxygen, 0.5bar at the exhaust. But:

• The expansion speed from the pre-chamber improves. Power from 820m/s vs 769m/s adds 14% pressure.
• The fuel is denser (EDA 900kg/m³, GUA allegedly 1550kg/m³) and it needs less oxygen - both increase the pressures.

30% more pressure brings much more than the apparent 1s loss. Denser propellants make also lighter tanks.

This idea fits also the fuel-rich staged combustion (if any useful) and the full-flow staged combustion, in an other thread.

Marc Schaefer, aka Enthalpy

[Enthalpy]

Some press papers claim an oxygen and methane engine is under development that uses a full-flow staged combustion cycle as in
http://www.scienceforums.net/topic/81051-staged-combustion-rocket-engines/
and while this has been done for axygen and methane, and may perhaps work with a handful of amines
http://www.scienceforums.net/topic/83156-exotic-pumping-cycles-for-rocket-engines/?p=805383
also there and following messages
http://www.scienceforums.net/topic/82965-gas-generator-cycle-for-rocket-engines-variants/
I believe soots prevents methane in a fuel-rich pre-chamber.

A more exotic cycle would make sense, where methane follows an expansion cycle, oxygen a staged combustion, and each pumps itself:

 

I plan to consider coupled shafts later. Here uncoupled:

• Besides hydrogen, only methane fits. Ethane, propane, cyclopropane, spiropentane... stay liquid around 300K and a few 10bar. Hard-to-light fuels are excluded.
• Each turbine drives a pump for the same propellant. Leaks are less critical, seals are easier. Each turbopump is smaller.
• Not needing to pump the oxygen, the methane expansion cycle achieves a decent chamber pressure.
• As the methane side limits the chamber pressure, the oxygen turbine is cooler than in a usual oxygen-rich staged combustion.
• Two cycles must be started at the same time.

The methane engine developed by SpaceX uses probably this cycle, without coupling. "Lower turbine temperature", "easier seals", "both propellants gaseous", "no soot" - and misleading "full-flow staged combustion".

The hybrid cycle brings technological advantages, but how efficient is it? This depends much on arbitrary choices.

• 800K=527°C out of the cooling jacket, already a lot. Flowing down protects the chamber better than up. A strong engine needs several long chambers.
• This holds when the engine throttles to 60% thrust, so at 100%, methane exits the jacket at 515K.
• P/2 expansion to the turbine is best, the isentropic work to 455K is 2760J/mol.
• A turbine 79% efficient, pump 74%, injector 88%, and no loss elsewhere (which is unfair) leave 187bar in the chamber.

Taking the usual oxygen-rich staged combustion Rd-170 as a reference, which obtains 3307m/s gas speed from Rg-1 "kerosene" at 535bar/245bar/0.8bar:

• Bulkier methane in a fair extrapolation of this cycle would burn at 228bar and gain 8s, Pmdeta 3s, cyclopropane 10s.
• The present hybrid cycle at 187bar gains 4s only.
• Does some decent amine mix work in a recombination prechamber? This full-flow cycle does bring performance and the same advantages.

Some improvement paths:

• Shutting some chambers off to throttle would gain much.
• Or heat the methane by an exchanger at the preburner mainly. Heavy, but the main chambers are fewer and shorter, and a staged injection can regulate the methane temperature when throttling. Cyclopropane then?
• Couple the shafts.

While the hybrid methane cycle eases the turbopump seals, it gains only 1s over Pmdeta. I'd prefer the hard-to-light fuel.

Marc Schaefer, aka Enthalpy

[Enthalpy]

A cycle where the oxygen-rich preburner heats the fuel in an exchanger to turbine it too would look like this:

It needs several flame stages in the preburner to regulate all temperatures, possibly several stages at the exchanger. At least, throttling doesn't make the fuel hotter.

I've computed the achievable pressures by expanding the gases at constant Cp: the power differs by 2% only from Propep. I've kept RD-170's 535bar at the prechamber, almost as good as 700bar. 500°C at the turbine resulted from oxidation then; I've also checked at 650°C, where nickel alloys creep little enough. I've also compared with RD-170's known oxygen-rich staged combustion. The hypotheses are arbitrary but the comparison is fair.

Unzip the two-sheet xls:
HybridStagedExchanger.zip

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The heat exchanger gains little performance, some 3s I_sp, and only for methane supposed to withstand heat. Accepting 10g decomposition products over 300s in 0.16m3 exchanger and ducts at 535bar permits 710°C for methane, 390°C for cyclopropane and 320°C for spiropentane, according to
http://kinetics.nist.gov/kinetics/Detail?id=1956SHA/PAV811:1
http://kinetics.nist.gov/kinetics/Detail?id=1961FAL/HUN609-611:1
http://kinetics.nist.gov/kinetics/Detail?id=1972FLO/GIB548:1
and a low temperature can't pay for pumping the fuel to a higher pressure.




Coupling the turbopumps brings very little performance. Safer start maybe. But easier seals are more important.

Cyclopropane and spiropentane outperform methane as their density lends to a higher pressure, and they shrink the tanks too. The advantage would be bigger at a gas generator cycle or with electric pumps. Cyclopropane is mass-produced, spiropentane could probably be.

Methane with a heat exchanger gains only 3s over cyclopropane without. If a fuel-rich pre-burner accepts methane somehow, this comparison will hold. To my opinion, not worth an exchanger nor a second pre-burner.

I've included Pmdeta in the table because it's more common than rocket "kerosene", more efficient, more resistent to fire. Safe and more efficient fuels may be possible
http://www.chemicalforums.com/index.php?topic=79637.msg290422#msg290422

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The exchanger is as badly difficult as expected. Spark-gap machining and molybdenum, niobium or tantalum alloys may contribute to a solution. Also, a 3s better 450t first stage gains at its end only 2t, which the exchanger can't squander. I won't put time in it.

Would methane soot in a fuel-rich pre-burner? The chemical equilibrium tells yes, so kinematics decides. Maybe little methane can burn with enough oxygen, the combustion get enough time to end (still a small throughput), and only then be quenched quickly with abundent methane. Experiments must decide.

Marc Schaefer, aka Enthalpy

[Enthalpy]

SpaceX gave information about their future Raptor methane engine
https://en.wikipedia.org/wiki/Raptor_(rocket_engine_family)
only a true full-flow, including one fuel-rich pre-chamber, can achieve 300bar in the chamber.

I've seen no report about soot or not at the pre-chamber and turbine. Soot would make the reuse of the engine more difficult and reduce the efficiency. Thermodynamic equilibrium tells "soot", but maybe the pre-chamber can first have a balanced flame, then quench it at once with the methane, so soot has no time to form at moderate temperature between the pre-chamber and the chamber.

In case soot remains an issue, replacing the methane-rich pre-chamber by the recombination of an auxiliary propellant would keep much of Raptor's design and performance.



Hydrogen peroxide, hydrazines and methylamine are badly dangerous, but I've described amine mixes that recombine at a good temperature without soot (paperwork!):
http://www.scienceforums.net/topic/82965-gas-generator-cycle-for-rocket-engines-variants/ and followings
http://www.scienceforums.net/topic/83156-exotic-pumping-cycles-for-rocket-engines/ and followings
check for grey shaded tables.

Liquid amine mixes can be fed from their tank by helium, and solid ones recompose in their casing, with turbine inlet vanes controlling the pressure hence speed of recomposition.

Raptor's vacuum 3.5MN and Isp=382s with 360:100 mix ratio need 203kg/s methane subcooled to 447kg/m³. Liquid injectors to 300bar and a pump, 88% and 74% efficient, take 20.9MW at the shaft. Expanding from 88bar to 1.8bar at a 79% efficient nickel alloy turbine need 23kg/s of added amine mix. This flux, 2.4% of oxygen+methane mass, expands further to 0.4bar and 730m/s to provide roll control and add 17kN (0.5%). The composite Isp drops by 1.9%, but 0.5% more thrust and propellants improve a first stage, whose performance drops by about 1.5%.

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An other alternative may be an oxygen-rich staged cycle with a molybdenum alloy turbine to keep a good chamber pressure.

At a new engine design, I'd rather have cyclopropane or spiropentane in an oxygen-rich staged cycle, or to smash methane's performance, cool the engine with the oxygen, and burn ethylene or if possible bicyclobutane.

Marc Schaefer, aka Enthalpy

[Enthalpy]

When a separate flux of propellant(s) is turbined and dump, it's a gas generator flux. If the whole flux of a main propellants burns partially in a prechamber and is turbined upstream the main chamber, it's a staged combustion cycle. But when a separate flux of propellant(s) is turbined and dumped in the main chamber together with the main propellant fluxes, let's call it an integral gas generator.

What isn't stupid in my proposal?

• Pressure vessel(s) can store the auxiliary flux(es). Start and re-start by opening valves, especially if a catalyst or mutual contact ignite the auxiliary flux(es).
• The auxiliary flux remains hot as it enters the main chamber and can ignite the main propellants. Some propellant added locally can help. Permanent ignition can enable faster main propellants speed, hence a smaller main chamber.
• No propellant mass is wasted. Main 100bar from 210bar or 350bar pressure vessel outperform a gas generator cycle at 140bar, already with steel tanks, clearly with fibre tanks.
• The integral gas generator is less efficient than a staged combustion, but it's a lot simpler, especially with clean auxiliary propellants like hydrogen peroxide, maybe amines mixes.
• Turbine speeds are easier than with a gas generator cycle, and they fit the pumps better.
• Membranes in the pressure vessels let start the pumps in zero gravity. Thrust from the auxiliary flux(es) can suffice to push the main propellants in the tanks' outlet.
• The auxiliary flux(es) can also provide attitude and orbit control to an upper stage after the main engine stops.

84% to 88% peroxide over a catalyst provide a reliable temperature that fits turbine alloys. 70% peroxide can't detonate but needs a fuel supplement for temperature. Maybe an amine, possibly unsaturated, brings hypergolic ignition. The amines mixes I propose elsewhere as alternatives to peroxide must first be tried: do they recompose where desired?

One turbine per main propellant would reduce the speed further. Injecting optionally some propellant between both turbines adjusts the mix ratio of the main propellants.

A sketch may come or not. The cycle is easy to imagine anyway.

Marc Schaefer, aka Enthalpy