[RobertGC]

A problem with scramjets, which attempt to achieve combustion at hypersonic speeds, is that the high speed means the air is producing so much drag that it is difficult to achieve net thrust when combusting with just the oxygen in the air.

So what if we also combusted with the nitrogen? Nitrogen makes up 80% of the air mass so perhaps this could provide sufficient thrust. The problem is nitrogen is commonly present in the atmosphere as N2, a molecule that is very stable, i.e., non-reactive.

But at hypersonic speeds so much heat is produced that the N2 is "cracked", generating various nitrogen species, such as N, NO, NO2, etc.

So how much energy could be produced by reacting hydrogen fuel with the various nitrogen species? How much for methane fuel?

The energy density and thus Isp would be less than for reacting with oxygen but conceivably the thrust could be greater because of the greater mass of the nitrogen.

[Borek]

And what would the "combustion with nitrogen" reaction be?

How much of the nitrogen is ionized/decomposed? Have you tried to check numbers, or is it just a wild speculation?

[Enthalpy]

Same opinion.

The formation of N₂ from two N atoms releases very much heat, so

• Combustion of hydrogen and methane occurs with the oxygen and leaves N₂ untouched. Forming NH₃ would release 45.9kJ/mol only, and don't expect it to happen at a few bar and 2000°C.
• Aluminium nitride would be less bad, releasing 318kJ/mol, but is quite impractical.
• Even when a spacecraft re-enters our atmosphere from low-Earth orbit, hence around 8km/s rather than targeted 2km/s hypersonic speed, N₂ doesn't atomize significantly. Only O₂ does, in small amount.
• But if using an amine as a fuel, the recombination of N atoms separated in the amine into N₂ is an essential contribution to the combustion heat.

"Achieving a net thrust" is difficult... for slightly different reasons. At 2km/s, the compression to subsonic speed prior to combustion heats the air to almost 2000°C, but the engine needs to expand hotter air so it's ejected faster than it came in and achieves a thrust. If you accept a dirty exhaust and you cool the engine properly, for instance aluminium burnt with the oxygen would bring the temperature increase. If you want a hydrocarbon, it's less easy, but its flame is hotter than in cold air.

The alternative known as scramjet seeks to burn the fuel in lukewarm air so the temperature increase is big. This implies a small compression, hence the air speed remains supersonic at the burner. The big difficulty is to stabilize the flame at supersonic speed (hence "supersonic combustion" ramjet), for which hydrogen fuel hopefully brings the necessary detonation speed.

Somewhere, I've proposed to create in a scramjet combustor local shock waves to permanently ignite a fraction of the fuel there before the flame propagates sidewards: the deflagration and detonation speeds wouldn't be a limit any more.

Still an other possibility has emerged with the Sabre engine
https://en.wikipedia.org/wiki/SABRE_(rocket_engine)
which evaporates the liquid hydrogen fuel to liquefy oxygen from the intake air for subsequent combustion. While I consider it pointless for space launch, it's a seducing idea for atmospheric flight.

[Enthalpy]

Out of curiosity, I've used Propep to estimate the compression from 2km/s and 55mbar, -60°C (20km altitude). At stagnation, the air would attain 220bar (wow!) and 1920K, so heating it further is difficult.

Flame      Expansion      Fuel
  K           m/s        (liquid except noted)
======================================================
3138          2846        Methane
3204          2864        RP-1 (~Diesel)
3180          2875        Diethylene Triamine
3183          2905        Diaminoguanidine (s) (much)
3269          2910        Cyclopropane
3318          2910        Bicyclo[1.1.0]butane
3285          2910        Diazetidycyclopropane
3340          2915        Cyclooctatetraene
3307          2923        Ethylene
3326          2936        Spiropentane
3329          2936        Di-spiropentyl (storable)
3394          2948        Cubane (s)
3224          3008        Nitromethane (much)
3215          3046        Hydrogen
3763          3213        LiAlH4 (s)
4733          3258        Aluminium (s) (much)
3982          3276        AlH3 (s)
4818          3357        Aluminium (molten) (much)
=====================================================

The ones meant to consume less oxygen (nitromethane, amines...) consume much fuel.

With 2km/s intake, expansion isn't much faster, so this parameter matters a lot. Methane is bad.

Cyclooctatetraene (COT) is the highest performing with limited drawbacks. "Only" carcinogen.
Di-spiropentyl and cubane must first be synthesized and mass-produced.
Among the easily flammable ones, ethylene is cheap and easier than hydrogen.
Hydrides consume less fuel than metals.

[Enthalpy]

Fuels that heat by recomposing may ease the supersonic combustion in a scramjet. After recomposition in a pre-chamber, they would be hotter than their ignition temperature when meeting the air, and hopefully light permanently, needing no flame propagation faster than the supersonic air.

Such fuels must avoid soot while recomposing, which imposes more hydrogen per carbon than most amines offers. I've given examples based mainly on the guanidine family and ethylene diamine, there
http://www.scienceforums.net/topic/83156-exotic-pumping-cycles-for-rocket-engines/#entry805383
and beginning there
http://www.scienceforums.net/topic/82965-gas-generator-cycle-for-rocket-engines-variants/
in the tables with light grey background.

The no-soot proportions depend on the pre-chamber pressure. This pressure can be attained by a gas in the strong tank or by a pump rotated by a turbine built between the recomposition chamber and the combustion chamber. The proposed mixes also limit to 700°C or 1100°C the temperature after recomposition to fit nickel alloys and prospectively molybdenum alloys; active cooling seems difficult here too.

Marc Schaefer, aka Enthalpy

[Enthalpy]

The unsaturated rocket fuels suggested there
http://www.chemicalforums.com/index.php?topic=91735.msg327577#msg327577
would be efficient as ramjet fuels too, here after air compression from 2000m/s:

Flame      Expansion      Fuel
  K           m/s
=====================================================
3289          2911        Divinylcyclobutane, cis-1,2
3340          2915        Cyclooctatetraene
3377          2952        [3]-syn-Ladderdiene
3367          2960        Spirohex-1-ene
3430          2989        Dispirooctadiene
=====================================================

Cis-1,2-divinylcyclobutane is as efficient as cyclooctatetraene (COT) and possibly cheaper to produce by dimerization of butadiene, see for instance patent US 3,594,434:
www.google.com/patents/US3594434 (click on Pdf for images)

[3]-syn-Ladderdiene and spirohex-1-ene outperform cyclooctatetraene and cubane, but their cost and ease of use are unclear, while dispirooctadiene may not even exist.

Marc Schaefer, aka Enthalpy

[Enthalpy]

Here's a sketch of the scramjet suggested on May 14, 2017 whose fuel recomposes to be hot at injection so the flame is (hopefully) stable in supersonic air. As already noted, soot-free recomposition is difficult, but some amines mixes should achieve it, according to Propep. The recomposition of methylamine is knowingly unstable but mixes producing higher temperatures should improve. Maybe I give a thought at the prechamber design some day.

The fuel injection pressure can be obtained from a gas in a strong fuel tank, not so heavy with graphite fibres, or from a rocket-style turbopump to achieve some 300bar. Both on the same sketch. Nickel alloys limit presently the gas temperature to 600-700°C before the turbine, more with cooled blades; maybe molybdenum or niobium alloys fit.

Reaching all the supersonic combustion zone with an injected gas looks uneasy, especially if pressure-fed. This improves with a convergent-divergent that results from the height and thickness distribution of fins that protrude in the air duct (only 8 drawn on the front view) and carry the fuel. The fuel has then a small air distance to cross, and the slope of the fin thickness is small even in a short engine. But such fins are more difficult to cool than an axisymmetric convergent-divergent. Intermediate designs are possible.

Marc Schaefer, aka Enthalpy

[Enthalpy]

This hydrogenolysis cycle reacts hydrogen with a hydrocarbon or amine to produce heat. The hot intermediate gas shall ignite spontaneously in a scramjet's supersonic air flow for a stable flame. The hot gas can also rotate a turbopump, useful since a pressurised hydrogen tank is heavy.

It resembles the cycle I described there for rockets
http://www.scienceforums.net/topic/83156-exotic-pumping-cycles-for-rocket-engines/
but using air instead of liquid oxygen.

The first hydrogenolysis prechamber uses more hydrogen to limit the gas temperature to ~600°C upstream a turbine of nickel alloy; niobium alloy would improve. Hydrogen can be injected in two stages, with a hotter first zone to speed up the hydrogenolysis, followed by a mitigation zone. Consider regenerative active cooling of the prechambers.

The optional second hydrogenolysis prechamber adds dense fuel to heat the intermediate gas further. More of the dense fuel also shrinks the tanks and improves the scramjet's performance but lets carry more fuel mass. No-sooting lets reach some 1100°C as the turbine exploits 100K to bring the fuels to potentially over 400bar so the intermediate gas has over 200bar at the main chamber injectors. Consider niobium and tantalum alloys, maybe molybdenum, according to the desired temperature.

Different dense fuels at the prechambers seem useless. Separate pumps for the different prechamber pressures, possibly in series, add complexity for a limited gain of the already good pressure at the main chamber injectors.

Expansion from 873K and 400bar to 200bar achieves some 1170m/s, of which a single-stage turbine of uncooled nickel alloy exploits most. Several stages gain little pressure at the main chamber injectors; they can power separate shafts and pumps. Suppression of the second prechamber would be the main advantage of active cooling or niobium turbine.

The hydrogen pump needs three centrifugal stages for 400bar, so consider trade-offs: a single stage leaves about 100bar at the main chamber injectors. Or try different pumps: a fast rotary screw or the hypothetic Pelton-Schaefer pump
http://www.scienceforums.net/topic/59792-convert-rocket-engines-to-hydrogen-or-methane/?do=findComment&comment=626301

All gases are reducing, but hydrogen may embrittle some alloys. Small leaks between reducing fluids aren't catastrophic. Hydrogen can cool parts of the engine.

This cycle lets use energetic fuels but the extra hydrogen prevents sooting at the main chamber too. Hydrocarbons should limit the nitrogen oxides in the exhaust. Here are some examples of easily produced fuels listed by their flash point. The proportion and temperature refer to the soot-limited hydrogenolysis, the speeds are 400bar -> 200bar and 130bar -> 100bar expansions from 873K.

  K   x:100   m/s   m/s
====================================================
1119   882   1000   624   Pmdeta
1296   446   1191   745   1,1-diazetidylcyclopropane
1239   368   1171   736   cis-1,2-divinylcyclobutane
1258   628   1135   708   azetidine
1256   430   1189   747   ethylene
====================================================

How fast is hydrogenolysis around 1200K, what catalysts to use if any? No idea. Opinions and suggestions welcome as usual!

==========

The fins in the previous message, and their combination with the injectors, fit also a flat engine, not only a round one.

Marc Schaefer, aka Enthalpy

[Enthalpy]

Ahum. Niobium and tantalum alloys are worse than molybdenum ones for a turbine - and molybdenum alloys haven't replaced nickel ones for reasons I ignore and can be excellent.

Tantalum alloys creep horribly at heat, and niobium alloys more. The least bad I found is the old T-222 (10% W, 2.5% Hf, 0.01% C in mass). It creeps at 0.1ppm/h under only 193MPa at 1093°C but weighs 16.7g/cm³. As opposed, the old molybdenum alloy TZM (0.5% Ti, 0.08% Zr, 0,02% C) creeps at 0.1ppm/h under 300MPa at 1100°C and weighs 10220kg/m³.

This lets a TZM disk of uniform thickness and 1100°C rotate at 265m/s tip speed, which improves if the centre is thicker and colder, as is usual for a turbine disk. A single stage of action turbine at 300m/s exploits nearly all the kinetic energy of a 700m/s gas and a good portion of 1100m/s.

Tantalum alloys would still be seducing for static parts of the engine if some added layer could protect them against the air at heat, but such a layer didn't pop up immediately on the Web, and some sources tell that none has been found. Absorbed gas make tantalum alloys brittle.

[wildfyr]

Enthalpy, if you don't mind me asking, what is your background? I see a great deal of fuel and aerospace discussion from you.

[Enthalpy]

My background isn't chemistry, as everyone noticed...

I've made electrical engineering with semiconductors and much electromagnetism, built a satellite, radars and sonars, mechanical engineering including facilities for crash-test.

Still no rocket fuels nor musical acoustics professionally up to now, but I'd be interested.

[Enthalpy]

The attached diagram gives figures on the hydrogenolysis pumping cycle. They are not optimum, just an example. Two 150kN engines power a small hypersonic airliner.

With a single-staged uncooled turbine of nickel alloy, and a single-staged hydrogen pump for simplicity , the cycle still achieves 100bar and 958K=685°C at the main chamber injectors, expanding to 1569m/s for good stirring with air and permanent lighting. The turbopump is as small as a car turbocharger. Alas, it probably needs a gear, maybe hydraulic.

The turbopump can also provide electricity and air to the plane. This combines nicely with the engine starter. I describe quick electric machines there
http://www.scienceforums.net/topic/73798-quick-electric-machines/?do=findComment&comment=737931

Myrcene C=CC(=C)CCC=C(C)C is a by-product of the paper industry. Nearly as efficient as divinylcyclobutane but with +44°C flash point. A eutectic with pinene would be nice, or with sesquiterpenes for the flash point - mind the pour point for aircraft flying at -70°C.

This cycle burns less hydrogen and more of the dense fuel, which reduces the tanks. I stayed 20% off the soot limit. It can burn more hydrogen to increase the range, then with less pressure and heat at the injectors, or with several stages at the hydrogen pump.

I've split the hydrogen injection at the first pre-chamber, in case hotter hence faster hydrogenolysis is wanted. The optional second pre-chamber provides hotter, denser fuel. If expanding at once from 114 to 100bar, the gas would attain only 493m/s.

Moderate ram compression at the main chamber lets use nickel alloys with little air cooling, as at the prechambers and ducts. The combustor and a part of the divergent are built like a rocket chamber, with a cooling jacket where the hydrogen flows between the pump and the pre-chamber (not detailed on the diagram). Hotter than 2073K is easy. Combustion at uniform pressure is suboptimum.

==========

Scramjet designs adapt to the speed by changing their dimensions and can become ramjets for smaller Mach number by changing their shape - nothing special with the present pumping cycles.

The hydrogenolysis cycle provides insufficient thrust at zero speed, but it is also a means to feed a turbojet with a hydrogen fraction.

==========

Ahum again. As it looks, everyone would like molybdenum alloys for turbines, but up to now they are brittle.

Marc Schaefer, aka Enthalpy

[Enthalpy]

The staged combustion cycle is well known at rocket engines and may well have been considered to feed a scramjet, ramjet, turbojet or turbofan - more probably than the recomposition and hydrogenolysis cycles which I hadn't read of.

It burns hydrogen in a pre-chamber with little oxidizer for a gas temperature bearable by the turbine and completes the combustion in a main chamber. The power can deliver bulky hydrogen at 190bar, but a single-stage uncooled turbine achieves rather 100bar as suggested there
http://forum.nasaspaceflight.com/index.php?topic=26952.msg825332#msg825332
Run turbofans, turbojets, ramjets and scramjets on hydrogen for range.

SpaceX try to use methane in their full-flow Raptor engine. Equilibrium software tells "soot", but methane is slow to decompose if not too hot, so maybe the nearly-stoechiometric flame in the prechamber can be quenched swiftly enough to avoid soot. Or they have an other trick. A heat exchanger separating the main methane from the flame would be heavy at a rocket, but reasonable at a plane.

Staged injection of hydrogen is desired in the first prechamber, but then ignition and reaction speed are no worry. Control the oxidizer amounts tightly. A second prechamber can reheat the fuel gas before injection in the main chamber, up to really hot if cooling the pipes and injectors actively. This shall ignite permanently the fuel in a scramjet main chamber.

Several oxidizers fit:

• Liquid oxygen from a tank, pressurized or pumped, a bit heavy;
• Liquefied air or oxygen from the atmosphere;
• Pressurized gaseous air from the atmosphere;
• But other sources like hydrogen peroxide, slurry of ammonitrate, trinitramide etc are heavy.

Lace, Sabre and Scimitar engines already developed the liquefaction or compression of air subcooled by hydrogen
https://en.wikipedia.org/wiki/Liquid_air_cycle_engine
https://en.wikipedia.org/wiki/SABRE_(rocket_engine)
https://en.wikipedia.org/wiki/Reaction_Engines_Scimitar
This work can be recycled here. No interesting thrust at zero speed from a scramjet, but the air throughput to the pre-chambers is much smaller, making the heat exchanger much easier.

Many comments on the recomposition and hydrogenolysis cycles apply here too.

Marc Schaefer, aka Enthalpy

[Enthalpy]

Hypersonic missiles are fashionable these days. Mach 6 has already flown years ago and is a reasonable limit to supersonic combustion (scramjet). Though, France lets MDBA develop an air-breathing missile for Mach 7 or 8
https://www.defensenews.com/global/europe/2014/11/29/france-studies-nuclear-missile-replacement/
which immediately raises the interrogation of a fuel efficient enough.

This French attempt seems to run into trouble, with question marks about the missile's mass compared with the aeroplane's capability, and whether stealth shall be abandoned to keep the speed.

Which would very well explain the query for more efficient fuels, and the search for help in Internet forums. After all, questioning the forums at the beginning of a project is common practice everywhere presently. Here it would have been delayed because the project is classified.

And I'm not convinced at all that MDBA would ask for help directly. Rather passing by an intelligence service that camouflages the question under a bizarre quest.