Hydralox with Oxygen Thrust-Augmented Nozzle?

[acksed]

We mostly know that the problem with hydrogen is the low density, requiring large tanks. Liquid oxygen is denser by far. It's even denser than kerosene or liquid methane, to the point that Starship's proposed propellant transfer is stated to be mostly LOX.

Compromises between tank size, nozzle size and thrust, among others, requires a more oxidiser-rich mixture ratio - I think around 5.5 to 6 O:F - when the greatest specific impulse of hydralox is found when it's more fuel-rich and with a vacuum-optimised nozzle.

Here's the proposal: a thrust-augmented nozzle can be over-expanded at sea level as long as fuel is injected downstream of the combustion chamber and the resultant pressure is equal or higher than atmospheric.

Injecting oxygen into a more fuel-rich hydrolox combustion stream, say 4 or even 3.5, to create a normal or even leaner O:F ratio of 6, 7, even 8 inside the nozzle would increase the combustion temperature, the thrust and maybe even the thrust-to-weight ratio of the whole rocket in the boost phase, as losing more oxygen at the start would reduce its mass more quickly.

I'm envisioning a sort of 1.5 stage to orbit, except with smaller drop tanks of oxygen.

A higher nozzle temperature would require more cooling, but also supply more energy to run a H2 expander cycle. (Yes, I know about the size limitations.) Oxygen cooling might be necessary on top of that, and I'm not sure you could run the pumps for the LOX off expander cycle, but Launcher's E-2 engine uses O2 for cooling at least.

Nozzle design would be tricky, no lie. It'd have to be large enough to provide good vacuum performance, small enough to fit into the rocket, resistant to extra hot oxygen... maybe mix oxygen and hydrogen cooling... and have injectors for the hypersonic nozzle flow.

Could it be simpler to design an engine to accept a greater range of mixture ratios? Yes. Could an extendable skirt do the same job? Yes... but they generally don't come back. This is a way to have a vacuum nozzle from the ground up, so to speak.

So has this ever been tried? Am I off-base entirely or is there something there?

[Robotbeat]

Yes. I like this idea.

[soyuzu]

Variable mixture ratio Hydrolox was a pretty common SSTO concept, I think.

Always liked the idea of dual mixture ratio hydrolox for any one crazy enough to try reusable SSTO. Start off running LOX rich, switch to fuel rich at the equivalent of staging. Bulk density similar to methlox, but with full flow staged combustion ISP of 360 at sea level LOX rich going to 460 fuel rich ISP in the vacuum. LOX rich tan mentioned by JonGoff on the other thread is an interesting variation, lower pressure injection of the extra TAN LOX reducing the delta in pumping requirements between fuel rich and LOX rich modes.  Would still pose some interesting pump design problems.

Actually you can make it very oxygen rich at the beginning, even to 12:1, to reduce chamber temperature. Basically a tri-propellant engine with two propellant.

[Robotbeat]

12:1?
“You are like a little baby!”
I’d go 26:1 for O:F ratio for a first stage for hydrolox. You want super high bulk density and low Isp in the beginning of flight, in this case 287s Isp. 26:1 O:F would give you bulk density nearly that of methalox.

[Bob Shaw]

What sort of engine materials could withstand white-hot O2 going through and past them? Would ablative nozzles be the answer?

[Robotbeat]

For an RLV, there’s another great argument for being oxygen-rich:

liquid oxygen is like 2 orders of magnitude cheaper than liquid hydrogen.

1 ton of LOx is about $100. 1 ton of LH2 is about $10k, and an absolute pain in the rear to transport and store. So the headache of using hydrolox can be reduced somewhat by being extremely LOx-rich in the first stage (or earlier part of the flight if you’re doing SSTO or whatever like in this thread).

With a 5.5:1 O:F ratio, 1000 tons of propellant costs $1.6M. For a 26:1 ratio, 1000 tons of propellant costs $470,000, a factor of 3.5 less. Tank and engine mass scale typically more closely to propellant volume than to mass, so if we assume the dry stage mass is 0.05kg per liter of bulk propellant, then for a first stage with a sea level Isp of 2.8km/s for a 26:1 O:F and 3.6km/s for a 5.5:1 O:F, and a required stage delta-v of 3.5km/s, then although the lower Isp option requires 39% more propellant mass, it requires about 35% LESS dry mass (which translates to lower fabrication costs… also total volume is lower just due to how much denser oxygen is than hydrogen) and its propellant costs are a factor of 2.5x lower.

So there’s pretty good argument in favor of LOx-rich lower altitude or first stage flight.

[acksed]

12:1?
“You are like a little baby!”
I’d go 26:1 for O:F ratio for a first stage for hydrolox. You want super high bulk density and low Isp in the beginning of flight, in this case 287s Isp. 26:1 O:F would give you bulk density nearly that of methalox.

You can do that? It might be like an oxygen-rich flamethrower at the start.

What sort of engine materials could withstand white-hot O2 going through and past them? Would ablative nozzles be the answer?

I think ablative nozzles are a single-use thing. You might need to line the nozzle with whatever SpaceX uses in their Raptor turbopumps, though I think the lower density of the exhaust and maybe some hydrogen curtain cooling could compensate.

[Robotbeat]

If you add *enough* oxygen, it actually reduces the temperature enough that material requirements aren’t as insane. Moderately oxygen-rich might be the worst place to be in terms of material requirements as the temperatures are very high but you’re also very oxygen-rich.

[Jim]

though I think the lower density of the exhaust and maybe some hydrogen curtain cooling could compensate.

Might as well then use the standard 6:1 ratio

[Jim]

12:1?

I’d go 26:1 for O:F ratio for a first stage for hydrolox.

That doesn't burn.  Need higher than 4% H2

[Robotbeat]

12:1?

I’d go 26:1 for O:F ratio for a first stage for hydrolox.

That doesn't burn.  Need higher than 4% H2

Correct… EXCEPT: That’s molar concentration (roughly, volume). 4 moles of molecular hydrogen is 8 grams. 96 moles of oxygen is about 3072 grams. That would be a 384:1 O:F ratio *BY MASS*, so we’re still well within the combustibility limits (you can double check this by adding up the chemical energy of the hydrogen plus the thermal heat needed to turn liquid hydrogen and liquid oxygen into gaseous hydrogen oxygen and hydrogen, and then high temperature gaseous water vapor (steam) and gaseous oxygen… it checks out and there’s plenty of energy left over).

From: http://conference.ing.unipi.it/ichs2005/Papers/120001.pdf

(Which also shows that at high pressure, the lower limit is actually ~6% molar ratio of hydrogen in oxygen… but again the 26 O:F mass ratio implies about a 38% molar percentage of hydrogen to total propellant… stoichiometric would be an 8:1 O:F, or a 67% molar ratio of hydrogen to total propellant, and a standard 6:1 implies 72% molar ratio of hydrogen to total propellant.)

[acksed]

So it can work.

What kinds of chamber and nozzle pressures in this hypothetical engine are we talking here? Thy olde RL10 has a modest 42-44 bar (no idea about nozzle pressure), while Aerojet's tests just went up to 34.47 bar for the chamber and an estimated 13.79 bar for the TAN.

On the other end of the scale is the RD-701: up to 294 bar in boost mode and a still-high 124 bar in sustain. Nozzle pressures unknown.

Lower pressure and high thrust is a good deal, but high chamber pressures seem to help wring out the last drop of performance.

Where's the sweet spot? What can we get away with if we want a 1.5 stage rocket with hydrolox engines packing oxygen-boosted TAN?

Side note: I found this article detailing lightweight liquid hydrogen tanks for fuel-cell aircraft. 67kg tank carrying 150kg H2 sounds good, but the devil is in the details. Composite tanks need a jig; large composite tanks need equally large jigs, and a defect means you have to scrap the whole thing. It might not stand up to rocket flight conditions either. Still, I'm interested.

The company itself says its tanks are also suitable for methane or oxygen. We would need lightweight drop tanks...

[Jim]

So it can work.

No, it can burn.  It doesn't mean the idea is viable.  Afterburning in a nozzle doesn't really help.

[Proponent]

At very high O/F ratios, what would be the power source for the pumps?

[redneck]

Using Kero/LOX for a thrust augmented nozzle alleviates some of these concerns. There are some techniques to get a very low L* in the augmented attachment. High thrust at lift off with the augmentation reducing or eliminating the Pa/Pc losses. High expansion ratio later when the augmentation gets out of the way for a high expansion ratio nozzle.

For first stage use launching in atmosphere. In vacuum, augmentation is a net loss.

[acksed]

At very high O/F ratios, what would be the power source for the pumps?

I don't have good answers for that, all this is just speculation, but it depends on how large I want the engine to be.

If I wanted a thrust-augmented equivalent to RL10, the expander cycle would probably still suffice for the hydrogen side, but for the oxygen side an expander cycle might need need to flow oxygen down the bottom of the nozzle - not that oxygen's a particularly good working fluid, we're taking advantage of the sheer mass flow - and it might additionally take a burner with heat exchanger to augment the heat gathered.

[Robotbeat]

At very high O/F ratios, what would be the power source for the pumps?

Staged combustion would still allow high power, although perhaps not as high chamber pressure as, say, Raptor.

[laszlo]

Watching Jim and Robotbeat debate makes my inner nerd tingle. This is why I paid to join this website. Thanks guys.

[sevenperforce]

At very high O/F ratios, what would be the power source for the pumps?

Staged combustion would still allow high power, although perhaps not as high chamber pressure as, say, Raptor.

At launch, you could use an expander cycle for the hydrogen side and ORSC for the oxygen side, since the amount of LOX you'll be pushing is far higher than the amount of LH2. You could also supplement the ORSC with a LOX-expander boost pump. As the ascent continued, you could throttle down the oxygen-rich preburner gradually, until the LOX-expander boost pump ended up doing the majority of the work.

One neat idea I saw from an old Aerojet patent was to use a preburner without any turbine, simply as a way to extract more heat to operate the expander cycle via a heat exchanger inside the preburner (attached). In their concept it was a fuel-rich preburner operating a single-shaft LH2 expander cycle, but you could imagine a similar arrangement with an oxygen-rich preburner, a turbine, a heat exchanger, and a split LOX-expander. Initially you'd be routing most of your oxygen-rich gas through a turbine, but you'd use a valve to gradually put less through the turbine and more through the heat exchanger, so that the expander turbine power increases while the preburner turbine power decreases.

[Robotbeat]

A 26:1 O:F ratio would imply roughly 3 times as much pumping power for a given chamber pressure (with similar amount of fuel flow). So reduce the chamber pressure to a third, and you'd have sufficient power. Raptor has ~300 bar, Merlin has ~100 bar. So you'd be down near Merlin chamber pressure, maybe BE-4 chamber pressure. Still not terrible at all. But further optimization is likely possible.

[edzieba]

So it can work.

No, it can burn.  It doesn't mean the idea is viable.  Afterburning in a nozzle doesn't really help.

That's kind of the crux: how far can TAN be pushed in terms of thrust?
For the original TAN concept of taking an very overexpanded bell and compensating with some burning within the bell to prevent flow separation (and get a bit of bonus thrust) we can be pretty sure it works, as it has been subscale tested.
But if you want to take the same concept and generate the majority - or even an appreciable fraction - of your thrust in post-throat combustion, I'm not so sure. You don't get the benefit of the DeLaval nozzle to accelerate the combustion products, and your nozzle extension needs to be strengthened.

[Robotbeat]

The concept of adding more propellant doesn’t require any combustion in the nozzle at all, BTW. You could use nitrogen or Argon, or even powdered rock, and get much the same effect. Adding oxygen is good because it’s basically as cheap as nitrogen, is liquid, and you’re carrying liquid oxygen anyway, plus you can get more complete combustion (as rocket engines typically run fuel-rich).

[sevenperforce]

If you want to pass everything through the combustion chamber and just vary the O:F ratio then here's that design with a preburner that initially functions as a classic ORSC turbopump, but then can alter the LOX flow in an expander cycle to pull heat off of the preburner and thus increase specific impulse while decreasing LOX flow and thrust.

[acksed]

So it can work.

No, it can burn.  It doesn't mean the idea is viable.  Afterburning in a nozzle doesn't really help.

That's kind of the crux: how far can TAN be pushed in terms of thrust?
For the original TAN concept of taking an very overexpanded bell and compensating with some burning within the bell to prevent flow separation (and get a bit of bonus thrust) we can be pretty sure it works, as it has been subscale tested.
But if you want to take the same concept and generate the majority - or even an appreciable fraction - of your thrust in post-throat combustion, I'm not so sure. You don't get the benefit of the DeLaval nozzle to accelerate the combustion products, and your nozzle extension needs to be strengthened.

I went searching and actually found the paper (embedded below). It turns out the reason it was using the LANTR chamber, injectors and nozzle was it was actually spinning off from the oxygen injection part of LANTR; I accidentally recreated a knockoff 'vegan' LANTR.

Anyway, the physical tests used that, water-cooling the nozzle and chamber while pushing it to about a mean of 40% RP-1/LOX augmentation, with a peak of 77% and a mass-flow of 108%, but the authors did some calculations on two bigger engines: the NK-33 and an RS-68-equivalent, both enhanced with TANs.

Their takeaways were that the NK-33's TAN could be expanded to a ratio of 58:1 with 40% augmentation while keeping the nozzle exit pressure of 6 psi/0.41 bar. This gave an average gain of 4.5 in specific impulse (assuming 20% of its mission is sea-level and the rest in vacuum) and an increase in T:W ratio from 128:1 to 150:1.

The hydrolox RS-68-like could be pushed further, increasing its T/W from 46:1 to "greater than 60:1". The graph of impulse over thrust augmentation given shows an increase in nozzle ratio to about 50:1 enabled a thrust augmentation of about 80% and an averaged increase in Isp of 11 seconds. That's pretty damn good.

Pushing the TAN to 3 times the mass-flow in just oxygen might be too much of an ask, though.

[InterestedEngineer]

So it can work.

No, it can burn.  It doesn't mean the idea is viable.  Afterburning in a nozzle doesn't really help.

That's kind of the crux: how far can TAN be pushed in terms of thrust?
For the original TAN concept of taking an very overexpanded bell and compensating with some burning within the bell to prevent flow separation (and get a bit of bonus thrust) we can be pretty sure it works, as it has been subscale tested.
But if you want to take the same concept and generate the majority - or even an appreciable fraction - of your thrust in post-throat combustion, I'm not so sure. You don't get the benefit of the DeLaval nozzle to accelerate the combustion products, and your nozzle extension needs to be strengthened.

I went searching and actually found the paper (embedded below). It turns out the reason it was using the LANTR chamber, injectors and nozzle was it was actually spinning off from the oxygen injection part of LANTR; I accidentally recreated a knockoff 'vegan' LANTR.

Anyway, the physical tests used that, water-cooling the nozzle and chamber while pushing it to about a mean of 40% RP-1/LOX augmentation, with a peak of 77% and a mass-flow of 108%, but the authors did some calculations on two bigger engines: the NK-33 and an RS-68-equivalent, both enhanced with TANs.

Their takeaways were that the NK-33's TAN could be expanded to a ratio of 58:1 with 40% augmentation while keeping the nozzle exit pressure of 6 psi/0.41 bar. This gave an average gain of 4.5 in specific impulse (assuming 20% of its mission is sea-level and the rest in vacuum) and an increase in T:W ratio from 128:1 to 150:1.

The hydrolox RS-68-like could be pushed further, increasing its T/W from 46:1 to "greater than 60:1". The graph of impulse over thrust augmentation given shows an increase in nozzle ratio to about 50:1 enabled a thrust augmentation of about 80% and an averaged increase in Isp of 11 seconds. That's pretty damn good.

Pushing the TAN to 3 times the mass-flow in just oxygen might be too much of an ask, though.

The more important metric than T/W is thrust per unit area.

Notice how the bottom of Booster is crammed with Raptor-2s, or SLS requires strapon boosters to get off the ground.  Any decrease in thrust per unit area will decrease the T/W of the entire rocket at the pad.

A large engine bell is contraindicated for the first 4-5 km of flight.  That 58:1 expansion ratio is going to require a lot of surface area. (Raptor-2 is 40:1)

[Robotbeat]

For Starship, maybe, but there are other concepts which go wider instead of taller.

[InterestedEngineer]

For Starship, maybe, but there are other concepts which go wider instead of taller.

how wide can you go?  (besides in KSP).

At some point air resistance becomes a problem, or building a wide enough launch mount (lots of people choked about Starship's 9 meter diameter, which can't be transported overland very far)

[Robotbeat]

It’s best to think of weight per unit cross section area. Starship is actually very tall and heavy for its cross section, but you don’t have to go that way. Starship gets extremely low drag because it’s only 9m wide but weighs like 5000tonnes. N-1 weighed about half that but was 17m wide. So in total it had a weight per area a factor of 6.5x smaller than that of Starship, plenty of room to reduce the thrust-per-area of the engine from Raptor’s value.

[InterestedEngineer]

It’s best to think of weight per unit cross section area. Starship is actually very tall and heavy for its cross section, but you don’t have to go that way. Starship gets extremely low drag because it’s only 9m wide but weighs like 5000tonnes. N-1 weighed about half that but was 17m wide. So in total it had a weight per area a factor of 6.5x smaller than that of Starship, plenty of room to reduce the thrust-per-area of the engine from Raptor’s value.

And now you have longer feed pipes.

Longer pipes can sympathetically vibrate with a lot more frequencies than short pipes.

Not sure how N-1 did with mass ratio, larger is more mass too.

[sevenperforce]

It’s best to think of weight per unit cross section area. Starship is actually very tall and heavy for its cross section, but you don’t have to go that way. Starship gets extremely low drag because it’s only 9m wide but weighs like 5000tonnes. N-1 weighed about half that but was 17m wide. So in total it had a weight per area a factor of 6.5x smaller than that of Starship, plenty of room to reduce the thrust-per-area of the engine from Raptor’s value.

And now you have longer feed pipes.

Longer pipes can sympathetically vibrate with a lot more frequencies than short pipes.

Not sure how N-1 did with mass ratio, larger is more mass too.

The N1's first stage had a dry mass of ~130 tonnes and carried ~1750 tonnes of LOX and RP-1, a fuel fraction of 93% (although 15 of those tonnes were expended to boiloff and throttle-up before the rocket even left the pad). With 3400 tonnes of propellant on board, Superheavy would need to be under 240 tonnes dry in order to improve on the N1's mass ratio. Of course, RP-1 is much more dense than CH4, so Superheavy is playing catch-up from the beginning.

The Saturn V's S-IC first stage, with a common bulkhead and monocoque tanks, was significantly better than the N1, at a fuel fraction of 95.4%.

[Proponent]

The Saturn V's upper stages definitely had common bulkheads, but I don't believe the S-IC did (see 61st page of the SA-507 Flight Manual, attached).

There were upgrade proposals in which the forward dome of the RP-1 tank would be replaced by honeycomb of mini domes to reduce the amount of wasted space.

[Proponent]

I’d go 26:1 for O:F ratio for a first stage for hydrolox. You want super high bulk density and low Isp in the beginning of flight, in this case 287s Isp. 26:1 O:F would give you bulk density nearly that of methalox.

Maybe there are other reasons for going that high, like reducing temperature or increasing thrust, but otherwise I would think you wouldn't want to go any leaner than the mixture ratio that maximizes impulse density (i.e., the product of specific impulse and bulk density).  That ratio will of course depend on chamber pressure and expansion ratio, but when I was playing around with a simple variable-mixture-ratio hydrolox SSTO model a few years ago, it came out at about 17.

EDIT: Corrected link to old post; now shows how O/F changes from lift-off to shutdown..

[Robotbeat]

Again, the reason I was interested in it was long term reducing the energy and/or propellant costs. It can make a big difference for the first stage.

Keep in mind that if you neglect stage dry masses (and gravity and drag) and consider just energy put into the reaction mass, an exhaust velocity equal to 62.75% of the mission delta-v maximizes the conversion efficiency (64.76% in this case) of reaction mass energy to payload kinetic energy IF you pick a constant exhaust velocity. IF you let the exhaust velocity change over time, the optimum exhaust velocity is whatever your current velocity is relative to your starting point (except for right at first, as you want to keep your total propellant mass finite LOL), and the efficiency can then approach 100% (assuming basically an infinitely long insulated nozzle extension… and perfect combustion efficiency…).

Maybe we’re all just condensed combustion exhaust from an enormous 4 dimensional rocketship, expanding since the Big Bang haha…

[1]

I went searching and actually found the paper (embedded below).

A key feature of this paper, assuming I'm reading it correctly, is this line:

These analyzes indicate that providing the LOX and hydrocarbon fuels to the TAN injectors from the boost pump discharge is feasible with minimal impact on the main turbopump assembly (TPA), preburners, and main chamber operating performance.

IMO, this suggests that the overall O/F ratio if the engine as a whole does not appreciably change. To me, this makes sense, as drastically increasing passive reaction mass comes with an opportunity cost in total available energy. If an increase in energy efficiency does not overcome a reduction in total available energy, then it's a poor trade. Better to just burn the less efficient engine for a longer time.

Pushing O/F more towards 'O' may have uses, but I wouldn't drift too far away from stoichiometric in that direction, either.
26:1 to me feels waaaaay to high. But maybe 10:1 early during boost, assuming the nozzle can survive? Might be reasonable. Haven't run any numbers on my end.

[Robotbeat]

Thing is that passive propellant like oxygen is virtually free, energy-wise, compared to fuels. For a given fuel flow (power input), you generate way more thrust by adding a huge amount of passive propellant.

It‘s like a jet engine with a larger bypass ratio, except the bypass air is oxygen that you’re carrying with you. A jet engine with a high bypass ratio is more energy-efficient than a turbojet. But like that analogy, it’s not advisable to use this configuration at high speeds because it’s no longer energy efficient to have accelerated all that reaction mass with you. There’s an optimum mission velocity for this sort of thing. Early in flight, it’s optimum to do this. Later in flight, it’s more efficient to operate at higher exhaust velocity (even though the thrust per unit power will be lower at higher exhaust velocity), ie without thrust augmentation by passive propellant but instead nearer to stoichiometric.

[InterestedEngineer]

Thing is that passive propellant like oxygen is virtually free, energy-wise, compared to fuels. For a given fuel flow (power input), you generate way more thrust by adding a huge amount of passive propellant.

It‘s like a jet engine with a larger bypass ratio, except the bypass air is oxygen that you’re carrying with you. A jet engine with a high bypass ratio is more energy-efficient than a turbojet. But like that analogy, it’s not advisable to use this configuration at high speeds because it’s no longer energy efficient to have accelerated all that reaction mass with you. There’s an optimum mission velocity for this sort of thing. Early in flight, it’s optimum to do this. Later in flight, it’s more efficient to operate at higher exhaust velocity (even though the thrust per unit power will be lower at higher exhaust velocity), ie without thrust augmentation by passive propellant but instead nearer to stoichiometric.

I'm having a tough time thinking of rocket oxidizer as "passive"

Helium?  Sure.

Hydrogen?  Mostly, save for a bit of embrittlement.

CH4?   Sure, minimal coking. Of course one can run fuel rich on a Raptor so it's not an `advanced topic`.

Oxygen?  Hot oxygen burns nearly everything except for the magic rare fairly low temperature high density compounds in the Oxygen path of Raptor.

why not just go straight for Krypton if mass is what is needed.  It's pretty cheap, $1.40 per kg


Cool pricelist btw:   http://www.leonland.de/elements_by_price/en/list

[1]

It‘s like a jet engine with a larger bypass ratio, except the bypass air is oxygen that you’re carrying with you.

Yes, but this is a critical difference. The jet engine actually gets its oxygen/air for free, so it's free to haul ONLY fuel. The rocket doesn't have this option. It must either haul both, or else it hauls dead/passive mass.

Here's what I believe to be the point of contention:

There’s an optimum mission velocity for this sort of thing.

Be very careful here. I agree that there's an optimal distribution for a fixed amount of energy available within a vehicle, but I caution against using 'exhaust velocity as close to zero as possible' a your preferred optimization point. If changing reference frames suddenly makes it appear that efficiency has dropped because the exhaust velocity is no longer zero, then your thought experiment is in danger of leading you astray in the general sense. Kinetic energy, after all, is not a conserved quantity.

Let's change the thought experiment somewhat. Instead of strap-on oxygen tanks, imagine that you have strap-on potassium tanks (while you're at it, also imagine that the potassium somehow behaves like a fluid). Now what? Well, first, a similar result should occur in as far as you're injecting a higher density substance into the exhaust stream, trading lower ISP for higher impulse density.

But, now that elemental potassium is free to react very violently with both the fully combusted hydrolox (water) as well as any other free oxygen in the exhaust.  So, by reacting with the combusted exhaust products, you'll increase the total energy of the exhaust in the nozzle. Will this therefore increase the pressure in the nozzle and therefore thrust of the vehicle? It might, especially early in flight and if the engine exhaust is slightly overexpanded.

By doing the 2K + H₂O -> 2KOH + H₂ reaction, more of your total energy may get bound back into that free hydrogen. Will that also now increase your ISP? If so, then this may be less "efficient" from an exhaust velocity standpoint. But if your total available energy increases, and if a portion of that increased energy goes into your vehicle and allows it to get better performance as a result, then we might not care.

Hope that makes sense somewhat. Overall I agree with you that there's likely an interesting trade to be made; I just don't believe going super oxy-rich is optimal all things considered.

[Robotbeat]

It‘s like a jet engine with a larger bypass ratio, except the bypass air is oxygen that you’re carrying with you.

Yes, but this is a critical difference. The jet engine actually gets its oxygen/air for free, so it's free to haul ONLY fuel. The rocket doesn't have this option. It must either haul both, or else it hauls dead/passive mass. ...

So you'd think that'd be a big benefit for jet engines, but it's actually not. You DON'T get that for free, you have to slow down the air first, or rather bring it up to the speed of the aircraft. (...which is the same thing the rocket does, when you think about it!)

SCRAMjets can get away with not doing that as much, BUT the impulse you can get out of a given mass of air that you DON'T slow down first is much less than that which you do. (This is a sort of difficult point to intuit, I think, but it is absolutely true.) So the benefit is negligible.

It's just the pure gravity losses of the extra weight that you lose out on, but that's much less if you have a nice high-thrust first stage with high T/W ratio.

[Robotbeat]

...

Be very careful here. I agree that there's an optimal distribution for a fixed amount of energy available within a vehicle, but I caution against using 'exhaust velocity as close to zero as possible' a your preferred optimization point. If changing reference frames suddenly makes it appear that efficiency has dropped because the exhaust velocity is no longer zero, then your thought experiment is in danger of leading you astray in the general sense. Kinetic energy, after all, is not a conserved quantity. ...

I promise you the physics works on this. See Geoffrey Landis explain it at 18:20 here:

[redneck]

I seem to recall that much of the thrust gain from augmentation was in reducing the losses from over expansion. One cold flow experiment from years back was a small jet reverse flow into the over expanded nozzle. Kicked the flow over to the side and entrained ambient air into the nozzle to achieve an aerospike like effect in a conventional over expanded nozzle. Noticeable increase in thrust.  All low pressure stuff (<200 psi) with compressed air.

[InterestedEngineer]

I seem to recall that much of the thrust gain from augmentation was in reducing the losses from over expansion. One cold flow experiment from years back was a small jet reverse flow into the over expanded nozzle. Kicked the flow over to the side and entrained ambient air into the nozzle to achieve an aerospike like effect in a conventional over expanded nozzle. Noticeable increase in thrust.  All low pressure stuff (<200 psi) with compressed air.

I would love to see the paper on that.

[Robotbeat]

Try NTRS.nasa.gov

[redneck]

I seem to recall that much of the thrust gain from augmentation was in reducing the losses from over expansion. One cold flow experiment from years back was a small jet reverse flow into the over expanded nozzle. Kicked the flow over to the side and entrained ambient air into the nozzle to achieve an aerospike like effect in a conventional over expanded nozzle. Noticeable increase in thrust.  All low pressure stuff (<200 psi) with compressed air.

I would love to see the paper on that.

I never wrote it up. I thought at the time that one of the Mojave companies was going to try my notchbell compensating nozzle.  Later I thought I could get something done when my finances permitted. I had realized by that time that the start ups couldn't afford to take too many chances on an unproven concept from an unproven source (me). The idea was that I would fund the notchbell through proof of concept and then move on to this one.

The rig was a 250 gallon tank pressurized on shop air to almost 200 psi. Blow down through a 1" hose to a nozzle mounted to a scale. No DAQ other than eyeball and memory. 1/2" throat to 1&1/2" exit. When the exterior air 1/8" hit at the right angle, the scale would jump about 5 pounds depending on the current blow down pressure. Lot of fun with sucking threads back into the nozzles when things worked right.

I could recreate it in a few weeks if there were a reason to do so. I don't really expect that to happen.

[VSECOTSPE]

See section 2.6 on “Rocket Augmentation for Turboaccelerator” in paper below.  Dumps an oxidizer in the afterburner of a turbine engine.  Similar to this thread’s concept.

https://www.sto.nato.int/publications/STO%20Educational%20Notes/RTO-EN-AVT-150/EN-AVT-150-02.pdf

[redneck]

See section 2.6 on “Rocket Augmentation for Turboaccelerator” in paper below.  Dumps an oxidizer in the afterburner of a turbine engine.  Similar to this thread’s concept.

https://www.sto.nato.int/publications/STO%20Educational%20Notes/RTO-EN-AVT-150/EN-AVT-150-02.pdf

Thanks for the link. I used to be familiar with several of those concepts with others being new to me. I remember my conclusions and some ideas from then. I may do some posting on my views when time permits. Most of these are too expensive and complex compared to conventional rockets, though there are some possibilities there.

[Ithirahad]

Instead of strap-on oxygen tanks, imagine that you have strap-on potassium tanks (while you're at it, also imagine that the potassium somehow behaves like a fluid).

You don't really have to "imagine" that the potassium behaves like a fluid, you just have to add some sodium. If NaK can be sprayed through a repurposed car fuel injector, you can blow it out the sides of a rocket nozzle just fine.