Author Topic: Pumping Cycles for Rocket Engines  (Read 45666 times)

Offline Enthalpy

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Re: Pumping Cycles for Rocket Engines
« Reply #20 on: 11/03/2011 01:29 pm »
If seeking the same moderate pressure and performance, a hydrogen staged combustion cycle is simpler (drawing) 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.

More details there
http://saposjoint.net/Forum/viewtopic.php?f=66&t=2272&start=60#p34927

Marc Schaefer, aka Enthalpy
« Last Edit: 11/03/2011 01:30 pm by Enthalpy »

Offline john smith 19

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Re: Pumping Cycles for Rocket Engines
« Reply #21 on: 11/03/2011 05:37 pm »
If seeking the same moderate pressure and performance, a hydrogen staged combustion cycle is simpler (drawing) 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.

Marc Schaefer, aka Enthalpy
You might like to look up the NASA SP8000 series of rocket engine component design monographs. Most rocket engines *already* use a single turbine blade stage to drive the pump stage and generally speaking *every* pump needs 1 stage of pump blading *except Hydrogen due to its low density. The major exceptions I am aware of are the astonishing 6 stage drive turbine in the LOX pre pump to get LOX up to 400psi in the SSME.

You still seem to be equating "novel" with better.



MCT ITS BFR SS. The worlds first Methane fueled FFSC engined CFRP SS structure A380 sized aerospaceplane tail sitter capable of Earth & Mars atmospheric flight.First flight to Mars by end of 2022 TBC. T&C apply. Trust nothing. Run your own #s "Extraordinary claims require extraordinary proof" R. Simberg."Competitve" means cheaper cheap SCramjet proposed 1956. First +ve thrust 2004. US R&D spend to date > $10Bn. #deployed designs. Zero.

Offline Enthalpy

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Re: Pumping Cycles for Rocket Engines
« Reply #22 on: 11/04/2011 01:07 am »
No. Most hydrogen engines with a gas generator cycle have two-stage turbines: Vulcain, RS-68... This is due to the light molecular mass of the hot gas and the big pressure ratio.

And yes, I'm speaking about hydrogen engines, so any one using >150b has several pump stages, for instance 3 stages at the main hydrogen pump for the SSME, RD-0120...

Offline Enthalpy

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Re: Pumping Cycles for Rocket Engines
« Reply #23 on: 02/13/2012 02:10 am »
This already described rocket pumping cycle can use an oxidizer that stays liquid down to -107C
http://forum.nasaspaceflight.com/index.php?topic=26952.msg814547#msg814547
good for Mars, so a fuel of similar melting point is desirable. Simultaneously, a flash point above our Biosphere's temperatures is safer; this corresponds roughly to a boiling point above +200C. Such a liquid range would have more uses.

---------------------------------

One known compound, 2,4,6-dimethyl-dodecane, melts at -112C and boils at +237C, so its flash point could be near +85C. 2,4,6-dimethyl-tridecane melts at -102C. Alas, these oddities were produced just for measurements by small-scale methods like Grignard reagent, as far as I know - and I ignore much.

But I hope to see similar compounds easier to produce. When geminal methyl groups replace each methyl branch, the melting and boiling points most often improve, and alkylation can mass-produce the compound as illustrated.

Isobutane and Isobutene (2-methyl-propene) give 2,2,4-trimethyl-pentane, the iso-octane produced by refineries in acid reactors to blend gasoline and increase its octane rating.
http://en.wikipedia.org/wiki/Alkylation#Oil_refining
The step with Isobutene repeats here once, selectively at the alkane's only tertiary carbon.
By the same process, 1-Butene shall end the compound to lower the autoignition temperature, in contrast with iso-octane. 2,2,4,4,6-hexamethyl-decane shall result.

---------------------------------

Pristane, or 2,6,10,14-tetramethyl-pentadecane, melts at -100C and boils at +296C, but induces autoimmune diseases in mice. One CH2 longer, Phytane doesn't have this reputation, but is about as scarce and over-expensive.

Again, I hope geminal methyl groups keep the low melting point and ease production by alkylation as illustrated.

now Isohexene instead of Isobutene shall bring the geminal methyl groups on every fourth backbone carbon. Where a melting point over +296C isn't necessary, the molecule can be shorter than Pristane, as drawn here with 2,2,6,6,10,10-hexamethyl-tetradecane. The straight tail participates in the low melting point and esay ignition.

---------------------------------

Both examples support many variants of course.
- More steps with Isobutene or Isohexene give a higher boiling point. Iso-octane can also be bought as a reactant, and probably 2,2,4,4,6-pentamethyl-heptane as a by-product from the refineries' alkylation.
- The straight butyl tail is long enough for easy ignition. Other alkenes, maybe odd ones, may optimize the melting point.
- High octane would need a short tail, or one ending as C(C)C, where the dimethyl period should be broken to keep a low melting point.
- A CC(C)C- or CCC(C)C- head may improve the melting point over CC(C)(C)C-. Propane or n-Butane would replace Isobutane then, but would the lack of a tertiary carbon let more Isobutene, or worse Isohexene, react with itself?

---------------------------------

The process may need several pots at different temperatures, or if the intermediate compounds must be separated. But each reactor is productive.
Isobutane, Isobutene, 1-Butene are very cheap, as are Ethene, Propene, 1-Hexene, 1-Octene. 1-Pentene and 1-Heptene maybe slightly less so. Isohexene or 4-methyl-1-pentene is widely available.
The reactants are highly flammable, the catalysts corrosive, the products little known: hardly flammable, not volatile, not corrosive.

Martian spacecraft are a limited market, but such compounds can have more uses:
- Easily storable fuel for other spacecraft.
- Rocket fuel, similar to kerosene with a wider liquid range, affordable and abundent.
- Fluids for hydraulics, thermal management, lubrication.
- Any engine would be safer with a high flash point fuel. The long-tailed compounds, with low autoignition temperature, fit Diesel engines and gas turbines, while short-tailed ones fit petrol engines better.
- Alkylation products from a refinery cost more than gas, kerosene, Diesel oil, and there wouldn't be enough reactants to replace all these. But a low freezing point and a high flash point open special uses, like race cars, boats for the future Arctic routes...

Please remember I'm no chemist, so better advice is necessary before any attempt.
Marc Schaefer, aka Enthalpy

Offline john smith 19

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Re: Pumping Cycles for Rocket Engines
« Reply #24 on: 02/13/2012 07:37 am »

Please remember I'm no chemist, so better advice is necessary before any attempt.
Marc Schaefer, aka Enthalpy

You need to give an Isp for the propellant combination you are proposing. There is a NASA program that can calculate these given certain thermochemical properties or the Isp program (and it's graphical front ends). Both are downloadable from the net.

Again to have a *chance* of being considered for any rocket application in a business context it has to be *better* than the incumbent propellants. Cost and toxicity are *not* major issues.

Isp is, density impulse  is, long term stability is.

This has been mentioned before. Novelty <> better.
MCT ITS BFR SS. The worlds first Methane fueled FFSC engined CFRP SS structure A380 sized aerospaceplane tail sitter capable of Earth & Mars atmospheric flight.First flight to Mars by end of 2022 TBC. T&C apply. Trust nothing. Run your own #s "Extraordinary claims require extraordinary proof" R. Simberg."Competitve" means cheaper cheap SCramjet proposed 1956. First +ve thrust 2004. US R&D spend to date > $10Bn. #deployed designs. Zero.

Offline Enthalpy

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Re: Pumping Cycles for Rocket Engines
« Reply #25 on: 02/13/2012 04:35 pm »
Toxicity is a huge concern for averyone, that's why every agency has had research programmes for years trying to replace the hydrazine family. A good cycle using an alkane is hence better than one using hydrazine - please use the link already given.

A low freezing point is extremely useful for a Martian module, and this is the other aspect the fuels improve.

The fuels suggested are long alkanes, which alone answers the Isp, density, long-term stability questions.

Just one remark: I know that one Pdf book emphasizes propellant density, but this matters for missiles, far less for space launch. Launchers switch to hydrogen even for first stages now, beginning with Delta IV.

Beyond the stone-old Isp programmes "with a graphical interface" you may consider RPA ("Lite" is free):
http://www.propulsion-analysis.com/
faster, convenient, more capable.

Offline tnphysics

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Re: Pumping Cycles for Rocket Engines
« Reply #26 on: 02/23/2012 01:18 am »
The amines I cite are not hydrazine nor its parents, precisely because of safety.
I understand. However in *common* rocket engineering usage that is what is being referred to. It's something you might like to keep in mind. Like using PSIA or PSIG instead of bar, Pascals or other more rational units.

Quote
TEPA is a very safe amine, dirt-cheap and produced in huge amounts (by Huntsman and more) as well as triallylamine.


Divinylcyclobutane gives properties similar to the cyclopropyl-things. UV dimerization of butadiene produces it with good yield and looks scalable to mass-production; electricity for the lamps would amount to 300/t.

And, yes, heat of formation isn't the whole picture, of course - if it were, I'd have included things like acetylene.
You might like to consider Propyne, which is used as part of the industrial fuel "MAPP" gas. Nearly as simple as most HC's but the strained bond gives it a useful higher level of Isp.

Quote
I prefer the glow-plug I described here to a catalytic igniter. Starts quickly, not sensitive to contamination, less specialized on one propellant. And with the glow-plug, hypergolic ignition isn't necessary neither, as soon as one propellant produces heat by decomposition.

I gave a thought to hydrazine replacements (at Saposjoint.net > Science > Technology) but I believe it's a dead end. I prefer the MON decomposition cycle, or pressurized methane-oxygen. Somewhere at the other site, I describe a cryocooler to keep oxygen, methane and even hydrogen liquid indefinitely.

The cycle that decomposes MON runs oxygen-rich, as Soviet engines have over decades. The cycle that recomposes amines runs fuel-rich and I too see it as an advantage. So does the hydrogenolysis cycle and it is one efficient way to burn propellants with performance similar to methane, which would otherwise prefer an oxygen-rich staged combustion.

You need to understand something about the way rocket engineering is done.

Most of it is expendable. It's used once and thrown away. You get *one* shot to do it right and because the hardware is either in free space or small pieces at the bottom of the ocean it's *almost* impossible to establish *exactly* what went wrong.

This makes rocket engineers *very* cautious.

They like things *simple* and they like things with *long* histories. The fact the *existing* solution is very dangerous and expensive is effectively *irrelevant*. The dangers are known and the costs (in terms of staff training, special equipment etc) are known and probably written off.

I'm unaware of *any* rocket engine using a glow plug ignition system. The ones that need ignition uses spark plugs with a side propellant flow (called augmented spark ignitors). The rest are hypergolic.

Your preference is irrelevant unless you're building rocket engines or you've been hired to do so by someone else. If you are *then* they matter.

I think you may have confused "different" (or merely complex) with *better*.  So what if your amine fuels are safer to handle. They are unproven (look up "Technology readiness level"), add complexity and deliver poorer performance. These are what *really* concern the people who write cheques for this stuff.

They'd only be used if you could demonstrate they would drop into *existing* systems with no change and lower the overall weight *more* than enough to compensate for the loss in performance.

But you can't.

A safer, lower performance alternative to the amines already exists and has flight experience. It's called high test peroxide and peopel *could* use if they wanted to.

You can handle this information in several ways.

1) re-work your ideas to fit them better to what is actually wanted, possibly along the way accepting that this is not possible.
2) Give up the idea entirely for the time being and learn more about the subject.
3) Insist you're right, "People don't understand my vision," etc etc and generally become yet another internet crackpot, one "resource" that the internet has *never* been short of.

You might like to note I'm the only person who's bothered to respond to your multiple posts. It's not about how much effort you put in. It's about working out weather it was necessary to put *any* effort in in the first place. Solving a problem no one else feels *is* a problem given the number of *real* problems that exist is pretty pointless.

If you want to go with 1) I'll suggest you consider catalytic ignition.

Iridium (the active ingredient of the Shell 405 catalyst used in both MMH decomposition for the Shuttle APUs and the aerojet GO2/LH2 cat ignitor) seems an efficient low temperature catalyst but I don't if this is the *best* transition metal for the job or merely the first one that worked (quite a lot of rocket engineering developed on this basis under the pressure of the Cold War. It was good enough to get the job done at the time and no one ever bothered, or could afford to find anything better).

Is it good enough for your amines? Is there a better catalyst?

BTW the idea of cracking a fuel to a more reactive one was suggested (IIRC) by Bob Zubrin for a US Navy hypersonic missile project. The standard US navy fuel (JP7?) is used to cool the skin and is cracked into ethyne (acetylene) which was expected to burn fast enough to be usable in a scramjet.

I'll leave it there.

Just FYI I am not sure that a catalyst would be possible for these amines. In fact, I have NEVER heard of them decomposing with evolution of heat!

Offline Enthalpy

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Re: Pumping Cycles for Rocket Engines
« Reply #27 on: 02/23/2012 01:48 am »
Other people introduced this unnecessary story of catalyst, I didn't.

EDA, DETA, TETA, TEPA other many amines would recompose into N2, CH4 and often soot with heat production. This results from their heat of formation, and since the produced heat brings the gas to around 600C with EDA+Guanidine, at least this reaction will proceed.

Offline Enthalpy

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Re: Pumping Cycles for Rocket Engines
« Reply #28 on: 03/03/2013 08:26 pm »
Hotter gas at a turbine gives more power to the pumps, and more chamber pressure makes a rocket engine more efficient - as long as the turbine doesn't melt. Molybdenum knowingly resists heat better than nickel alloys common at turbines; here are performance estimates.

At the LME, molybdenum trades at 24 usd/t in November 2012, nickel at 16 usd/t. This is not the price for the usable alloy, though.

Plansee had a Pdf called "Material properties and applications" for Molybdenum, other providers exist
http://www.plansee.com/en/Materials-Molybdenum-402.htm
which gives for TZM alloy 10280 kg/m3 and 10-4 /h creep rate at 400 MPa and 1100 C, while nickel superalloys would achieve 700 C.

400 MPa isn't much, but a turbine full disk at 600 C could accept ~580 MPa to achieve 350 m/s blade speed. This is enough for a hot oxygen turbine - ask someone else about oxidation and coating...

I compare with nickel alloy in an RD-170 engine (of course!). Hot oxygen enters the turbine at 772 K: the same margin allows 1173 K for TZM. Being equally over-optimistic, expansion from 509 b to 245 b accelerates the gas to 518 m/s with Ni and 642 m/s with Mo - fully exploitable with 350 m/s blades.

1.54 times more power brings the main chamber pressure from 245 b to 376 b. Expanded to 0.8 b, this gains 8 s. Not huge, but Boctane costs its shiny penny to gain 4 s, and methane would bring 11 s with a deep engine redesign.

Marc Schaefer, aka Enthalpy

Offline Enthalpy

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Re: Pumping Cycles for Rocket Engines
« Reply #29 on: 03/03/2013 08:33 pm »
As an alternative to heavy pressurized tanks and to complicated turbopumps, an electric pump can feed the propellants in the chamber(s). Conceivable at small chemical thrusters, where a high pressure improves the efficiency and injects enthalpy from the Solar panels.

-----

Scale up: electronics can control 1MW motors from the main's voltage. A 70% efficient centrifugal pump brings then 72kg/s of oxygen and farnesane C15H32 to 96b; expansion from 80b to 0.02b in a 2.6m nozzle to push 260kN with Isp=375s, enough for the main engine of a 30t upper stage.

Rotating at 780Hz
in vacuum, the motor is small. Its rotor can be a permanent magnet of Magnetoflex 93, D=250mm L=200mm. The D=350mm 3-phase 8-pole stator loses <1kW in its braided "Litz" wires and 70W in the Nanoperm magnetic circuit; it's cooled by the fuel. The motor weighs 140kg. It accelerates in 30s.

The safer Li-MnO2 primary battery brings ~650kJ/kg in the too quick discharge. The stage bringing 5900m/s would burn 24t; at full thrust, this would require a 470kg battery (20kg per ton of propellants), but throttling is easy, useful, and lightens the battery.

The battery is easy to integrate and welcome for the gimbal actuators, the igniters...

-----

Pressurized steel tanks for the same 24t but with only 36b in the chamber, throttling to 20b, would weigh 1930kg and provide 13s less Isp. Graphite tanks could weigh about 1100kg.

A gas generator cycle could waste 42kg per ton of propellants, later ejected at 1500m/s thus counting as 35kg/t - but this equivalent overhead is ejected all the way and a battery supposedly not.

Staged combustion is more efficient.

-----

Scale up further: power components control 6MW railway engines. This would provide 1MN thrust with 120b in the chamber, enough for 115t at a second stage or 70t per 6MW slice at a first stage.

-----

Smaller roll and vernier engines can also have electric pumps, for instance at a solid engine stage.

For a lander or descent-ascent module, I like the ease of starting and restarting electric pumps.

-----

I plan to check what fuel cells achieve presently, for a hydrogen-oxygen engine with electric pumps.

Marc Schaefer, aka Enthalpy

Offline Tass

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Re: Pumping Cycles for Rocket Engines
« Reply #30 on: 03/03/2013 09:04 pm »
One oxidizer is Mon-33, or 33% NO dissolved in 67% N2O4. It freezes at -107C, so if paired for instance with 2,4,6-trimethyl-tridecane (freezes at -102C), they stay indefinitely on Mars or an asteroid or a Moon just in white tanks.

What does freezing point have to do with whether they will keep on Mars? You should quote vapor pressures not freezing points. 
« Last Edit: 03/03/2013 09:29 pm by Tass »

Offline Andy USA

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Re: Pumping Cycles for Rocket Engines
« Reply #31 on: 03/04/2013 12:02 am »
No one is allowed to shout down a thread, or they will have their posts removed. As I just did.

Offline Enthalpy

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Re: Pumping Cycles for Rocket Engines
« Reply #32 on: 03/04/2013 12:08 am »
Electric pumps, but for a hydrogen engine now: fuels cells are still far too heavy. Here an RL-10B equivalent with Li-MnO2 primary batteries instead:

Injectors shall drop 17% pressure to 43.6b and pumps be 65% efficient, then the shafts need 145kW and 399kW, and the motors cumulate 80kg. The battery weighs 36kg per ton of propellants, a bit less by throttling.

This is mass where not desired, and an expansion cycle performs better, but the electric pump is simpler.

Offline Jim

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Re: Pumping Cycles for Rocket Engines
« Reply #33 on: 03/04/2013 11:26 am »

This is mass where not desired, and an expansion cycle performs better, but the electric pump is simpler.

Wrong conclusion and unsupported by data.
Electric pump is more complex.  It requires pump, battery, charging system, control system, wiring, etc.   It has no advantages over the expander.
The expander system is self contained and self controlling.
It is more efficient weight wise, ISP wise and complexity wise.

The fact that nobody is trying it is a big clue and it isn't for lack of the idea.
« Last Edit: 03/04/2013 11:28 am by Jim »

Offline Jim

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Re: Pumping Cycles for Rocket Engines
« Reply #34 on: 03/04/2013 11:30 am »
As an alternative to heavy pressurized tanks and to complicated turbopumps,

False conclusions.
The high pressure tanks are not "heavy" and still lighter than battery and motor.  You make unsupported assumptions in your math.
« Last Edit: 03/04/2013 11:33 am by Jim »

Offline Enthalpy

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Re: Pumping Cycles for Rocket Engines
« Reply #35 on: 03/04/2013 04:22 pm »
My suggestion is fully supported by maths, of course and as always.

Progress does exist. Batteries and electric motors are wrongly perceived as heavy, which suffices to explain that no-one was trying.

Offline Jim

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Re: Pumping Cycles for Rocket Engines
« Reply #36 on: 03/04/2013 04:25 pm »
My suggestion is fully supported by maths, of course and as always.

Progress does exist. Batteries and electric motors are wrongly perceived as heavy, which suffices to explain that no-one was trying.

not in this case.  You have shown nothing of the sort.
This is not progress, just throwing ideas against the wall to see if something sticks.  In this case, you are even missing the wall.

And wrong, the you are making the wrong perception.  The propellant tanks for a pressure system are not heavy.  Also, you have not quantified an actual system weights.   You are only making unsupported projections of weights.
« Last Edit: 03/04/2013 04:27 pm by Jim »

Online Robotbeat

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Re: Pumping Cycles for Rocket Engines
« Reply #37 on: 03/04/2013 04:33 pm »
...
Progress does exist. Batteries and electric motors are wrongly perceived as heavy, which suffices to explain that no-one was trying.
No, they're still much heavier than a gas generator or the like for the same power.
Chris  Whoever loves correction loves knowledge, but he who hates reproof is stupid.

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Offline Enthalpy

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Re: Pumping Cycles for Rocket Engines
« Reply #38 on: 03/04/2013 05:04 pm »
The propellant tanks for a pressure system are not heavy.
This alone disqualifies your contribution.

Offline Enthalpy

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Re: Pumping Cycles for Rocket Engines
« Reply #39 on: 03/04/2013 05:11 pm »
[Batteries and motors] are still much heavier than a gas generator or the like for the same power.
Agreed with heavier, and that is what I've written - but not much heavier if you consider the amount of derived propellants, not just the hardware.

This unexpected result alone justifies the discussion. The electric option is light enough to make the method viable though not optimum, and its development is easier than a gas generator.

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