Dual monopropellant supersonic combustion rocket

[sevenperforce]

A cursory examination of rocket engineering rapidly identifies LH₂/LOX rocket engines as the gold standard for efficiency. It's hard to beat 450+ seconds of vacuum specific impulse, even though hydrolox rockets don't always have the greatest T/W ratios and require very large fuel tanks which eat into mass fraction rapidly.

I think I can beat that, though.

Hydrogen peroxide is not a very good monopropellant. It has only 161 s of impulse. Hydrazine is a bit better, as far as monopropellants go; it boasts upwards of 220 s. And together, they're not much improved; a hydrazine/peroxide bipropellant rocket can't even break 300 s in a vacuum. 

Put them together in the right way, though, and I think I might be on to something.

This is a fairly basic, no-nonsense linear aerospike engine. There's just a single difference: instead of using small bipropellant combustion chambers, it uses staggered monopropellant combustion chambers, alternating between hydrazine and high-test peroxide.

The peroxide thrusters and hydrazine thrusters produce flows which, after beginning to expand against the aerospike, are already moving very fast. High-test peroxide's 161 seconds of specific impulse corresponds to an exhaust velocity of around 1.58 km/s while hydrazine's 220 seconds corresponds to an exhaust velocity of around 2.16 km/s. At a molar mass ratio of 3:2 (peroxide:hydrazine), the mutual flow is traveling down the aerospike at an average velocity of 1.83 km/s.

But it doesn't stay that way. As the two compressed supersonic flows mix, they ignite with each other:

Decomposed hydrazine contains 4 grams of diatomic hydrogen per mole; decomposed peroxide contains 16 grams of diatomic oxygen per mole. At the previously-mentioned 3:2 molar ratio, the oxygen and hydrogen will burn with a vacuum specific impulse of 430-450 seconds. Of course, the reactants compose only one third of the mass of the flow, so the net increase in propellant flow speed will be about 2.49 km/s.

However, because that increase takes place in a flow which is already moving at 1.83 km/s, the speeds stack. This staged combustion results in a total exhaust velocity of 4.32 km/s, for a specific impulse of 441 seconds.

Because monopropellant thrusters are being used, the thrust-to-weight ratio will be fantastic, a major advantage over other linear aerospike designs. Moreover, both fuels are dense and liquid at room temperature, allowing small tank volume and a smaller launch vehicle.

Edited February 16, 2016 by sevenperforce

[Steel]

Interesting idea. I have a couple of feasibility questions though.

52 minutes ago, sevenperforce said:

Interesting Decomposed hydrazine contains 4 grams of diatomic hydrogen per mole; decomposed peroxide contains 16 grams of diatomic oxygen per mole. At the previously-mentioned 3:2 molar ratio, the oxygen and hydrogen will burn with a vacuum specific impulse of 430-450 seconds. Of course, the reactants compose only one third of the mass of the flow, so the net increase in propellant flow speed will be about 2.49 km/s.

However, because that increase takes place in a flow which is already moving at 1.83 km/s, the speeds stack. This staged combustion results in a total exhaust velocity of 4.32 km/s, for a specific impulse of 441 seconds.

What I'm thinking is as follows (and I may be wrong, if so please correct me):

• Just putting two streams of bi-propellant together and igniting them you get a reaction where the distribution of momentum is essentially random (i.e the reaction exhausts in all directions). This means there will be some exhaust travelling at 4.3 km s^-1, but equally there will be some travelling back against the flow at 0.7 km s^-1 ,thus disrupting the flow, and some traveling out sideways from the aerospike. If this is the case then you'll not gain any advantage.
• Just intersecting two flows of propellant does not mean they will all react, so there may be unburned propellant in the exhaust. 

Edited February 16, 2016 by Steel

[kunok]

You need the reaction in some kind of chamber, or you will lose most of the power because the gas expands in all the directions, not only to the aerospike.

That's not taking in to account that the chemical process you describes doesn't sound right at all for me, but that's not my field of knowledge

Edited February 16, 2016 by kunok

[fredinno]

1 hour ago, sevenperforce said:

A cursory examination of rocket engineering rapidly identifies LH₂/LOX rocket engines as the gold standard for efficiency. It's hard to beat 450+ seconds of vacuum specific impulse, even though hydrolox rockets don't always have the greatest T/W ratios and require very large fuel tanks which eat into mass fraction rapidly.

I think I can beat that, though.

Hydrogen peroxide is not a very good monopropellant. It has only 161 s of impulse. Hydrazine is a bit better, as far as monopropellants go; it boasts upwards of 220 s. And together, they're not much improved; a hydrazine/peroxide bipropellant rocket can't even break 300 s in a vacuum. 

Put them together in the right way, though, and I think I might be on to something.

This is a fairly basic, no-nonsense linear aerospike engine. There's just a single difference: instead of using small bipropellant combustion chambers, it uses staggered monopropellant combustion chambers, alternating between hydrazine and high-test peroxide.

The peroxide thrusters and hydrazine thrusters produce flows which, after beginning to expand against the aerospike, are already moving very fast. High-test peroxide's 161 seconds of specific impulse corresponds to an exhaust velocity of around 1.58 km/s while hydrazine's 220 seconds corresponds to an exhaust velocity of around 2.16 km/s. At a molar mass ratio of 3:2 (peroxide:hydrazine), the mutual flow is traveling down the aerospike at an average velocity of 1.83 km/s.

But it doesn't stay that way. As the two compressed supersonic flows mix, they ignite with each other:

Decomposed hydrazine contains 4 grams of diatomic hydrogen per mole; decomposed peroxide contains 16 grams of diatomic oxygen per mole. At the previously-mentioned 3:2 molar ratio, the oxygen and hydrogen will burn with a vacuum specific impulse of 430-450 seconds. Of course, the reactants compose only one third of the mass of the flow, so the net increase in propellant flow speed will be about 2.49 km/s.

However, because that increase takes place in a flow which is already moving at 1.83 km/s, the speeds stack. This staged combustion results in a total exhaust velocity of 4.32 km/s, for a specific impulse of 441 seconds.

Because monopropellant thrusters are being used, the thrust-to-weight ratio will be fantastic, a major advantage over other linear aerospike designs. Moreover, both fuels are dense and liquid at room temperature, allowing small tank volume and a smaller launch vehicle.

The toxicity of hydrazine is a major disadvantage.

[DerekL1963]

 

2 minutes ago, fredinno said:

The toxicity of hydrazine is a major disadvantage.

As is the difficulties of handling hydrogen peroxide.

[shynung]

24 minutes ago, kunok said:

You need the reaction in some kind of chamber, or you will lose most of the power because the gas expands in all the directions, not only to the aerospike.

In an aerospike design, the atmospheric pressure acts like half a nozzle bell, the other half provided by the 'spike'. This works out like a rocket nozzle, but is self-adjusting, so a traditional rocket which burns the fuel in a chamber has a relatively consistent specific impulse despite changing atmospheric pressure.

The thing is, this design reacts the propellant mix outside a combustion chamber, so it's probably more dependent on atmospheric pressure than an aerospike nozzle on a traditional rocket. This could mean that, rather than seeing a relatively constant (or rising) specific impulse we see in traditional rockets as atmospheric pressure drops, this design loses specific impulse as it rises because, again, there's only half a combustion chamber for them to react in.

[kunok]

1 minute ago, shynung said:

In an aerospike design, the atmospheric pressure acts like half a nozzle bell, the other half provided by the 'spike'. This works out like a rocket nozzle, but is self-adjusting, so a traditional rocket which burns the fuel in a chamber has a relatively consistent specific impulse despite changing atmospheric pressure.

The thing is, this design reacts the propellant mix outside a combustion chamber, so it's probably more dependent on atmospheric pressure than an aerospike nozzle on a traditional rocket. This could mean that, rather than seeing a relatively constant (or rising) specific impulse we see in traditional rockets as atmospheric pressure drops, this design loses specific impulse as it rises because, again, there's only half a combustion chamber for them to react in.

The aerospike replaces the nozzle bell, not the combustion chamber, in both you need the combustion chamber to direct the exhaust of the gases (among other things like better combustion efficiency), the aerospike or the bell are for expanding the gas, changing pressure for speed.

We are almost saying the same. I'm not sure what did you understood.

[shynung]

Just now, kunok said:

The aerospike replaces the nozzle bell, not the combustion chamber, in both you need the combustion chamber to direct the exhaust of the gases (among other things like better combustion efficiency), the aerospike or the bell are for expanding the gas, changing pressure for speed.

We are almost saying the same. I'm not sure what did you understood.

This design is equivalent to two combustion chambers, each fed separately with hydrazine and HTP, feeding a shared nozzle. The idea is that the products of the combustion chamber was to react in the nozzle, producing more energy.

Also, because of the aerospike nozzle, the combustion products essentially react in a nozzle in which half of it is comprised of the ambient atmosphere, the other half being the spike itself. This could work as a decent nozzle/chamber contraption, but its effectiveness decreases as ambient pressure falls, which means the secondary combustion's energy is captured less by the nozzle itself.

[DerekL1963]

17 minutes ago, shynung said:

This design is equivalent to two combustion chambers, each fed separately with hydrazine and HTP, feeding a shared nozzle. The idea is that the products of the combustion chamber was to react in the nozzle, producing more energy.

AIUI Generally you want combustion to occur before the point of highest density and velocity (usually the throat of the engine) for maximum efficiency.   While this engine does capture some energy that would otherwise be wasted, it goes out of it's way to create that wasted energy...  My gut (which has nothing really backing it up to be honest) says there's something basically wrong with this scheme.

[sevenperforce]

1 hour ago, Steel said:

Interesting idea. I have a couple of feasibility questions though.

What I'm thinking is as follows (and I may be wrong, if so please correct me):

• Just putting two streams of bi-propellant together and igniting them you get a reaction where the distribution of momentum is essentially random (i.e the reaction exhausts in all directions). This means there will be some exhaust travelling at 4.3 km s^-1, but equally there will be some travelling back against the flow at 0.7 km s^-1 ,thus disrupting the flow, and some traveling out sideways from the aerospike. If this is the case then you'll not gain any advantage.
• Just intersecting two flows of propellant does not mean they will all react, so there may be unburned propellant in the exhaust. 

Good questions.

1 hour ago, kunok said:

You need the reaction in some kind of chamber, or you will lose most of the power because the gas expands in all the directions, not only to the aerospike.

That's not taking in to account that the chemical process you describes doesn't sound right at all for me, but that's not my field of knowledge

As I understand the physics/chemistry, this is going to be a case of shockwave-induced supersonic ram combustion, which is almost impossibly difficult to achieve in scramjets (which must mix a supersonic airflow with a subsonic fuel flow), but will be comparatively straightforward here because both exhaust flows are supersonic. The difference in speed will result in a standing shockwave where they meet; careful shaping of this standing wave by adjustment of the thruster expansion ratios and alignment can be used to ensure complete combustion.

The shape of the aerospike might need to be slightly suboptimal in order to ensure the proper thrust vector, but this will only bleed a few percent of vacuum ISP at most. I'm about 98% sure that an aerospike surface can be used as a combustion chamber in supersonic flow conditions.

The flow pressure of the propellants serves to confine and direct the combustion. The monopropellants have different molar masses and speeds, meaning that their momentum can be configured to ensure proper combustion as well.

1 hour ago, fredinno said:

The toxicity of hydrazine is a major disadvantage.

But think of the children! 

You know, the ones who want to grow up to be astronauts.

1 hour ago, DerekL1963 said:

As is the difficulties of handling hydrogen peroxide.

We figured out how to handle liquid hydrogen; high-test peroxide presents different challenges but no less surmountable ones.

52 minutes ago, shynung said:

The thing is, this design reacts the propellant mix outside a combustion chamber, so it's probably more dependent on atmospheric pressure than an aerospike nozzle on a traditional rocket. This could mean that, rather than seeing a relatively constant (or rising) specific impulse we see in traditional rockets as atmospheric pressure drops, this design loses specific impulse as it rises because, again, there's only half a combustion chamber for them to react in.

Interesting. I wonder if gimballing the thrusters to direct them further in or further down could compensate for this.

It will have to be an extended aerospike regardless, in order to allow for combustion time, but the weight cost of the extended aerospike will be negligible compared to the dramatically higher T/W ratio.

And I had promised you those maths, so as soon as I can get Excel up and running with the optimization equations, I'll post them. Anyone happen to have a handy hydrolox specific-impulse-vs-ratio table?

7 minutes ago, DerekL1963 said:

AIUI Generally you want combustion to occur before the point of highest density and velocity (usually the throat of the engine) for maximum efficiency.   While this engine does capture some energy that would otherwise be wasted, it goes out of it's way to create that wasted energy...  My gut (which has nothing really backing it up to be honest) says there's something basically wrong with this scheme.

The key here is that the standing wave between the two supersonic flows re-creates a new density/velocity maximum which exists within an already-moving reference frame, allowing exhaust velocity to stack. At least, that's how I've designed it; I have no way of knowing whether it will actually work.

I've poked around looking for something like this and I've never found anything, so either it's an absolutely horrible idea which is prohibitively inefficient for reasons which currently escape me, or it's something that simply never occurred to anyone before.

(edit):

P.S. You can look at different scramjet designs for a better idea of how such nozzles end up being shaped. The converging/diverging nozzle is no longer optimal for supersonic flow. Scramjets don't work, of course, but this is different because both flows are supersonic because they're both already-expanded exhaust plumes.

Edited February 16, 2016 by sevenperforce
Adding info.

[shynung]

8 minutes ago, DerekL1963 said:

AIUI Generally you want combustion to occur before the point of highest density and velocity (usually the throat of the engine) for maximum efficiency.   While this engine does capture some energy that would otherwise be wasted, it goes out of it's way to create that wasted energy...  My gut (which has nothing really backing it up to be honest) says there's something basically wrong with this scheme.

This isn't the first time the idea of reacting propellants in the nozzle rather than in the chamber had arisen. Pratt & Whitney used it in their Triton bimodal nuclear thermal rocket design, though only because they don't want to spray the nuclear fuel rods with LOX so as to not destroy them.

[sevenperforce]

3 minutes ago, shynung said:

This isn't the first time the idea of reacting propellants in the nozzle rather than in the chamber had arisen. Pratt & Whitney used it in their Triton bimodal nuclear thermal rocket design, though only because they don't want to spray the nuclear fuel rods with LOX so as to not destroy them.

I love bimodal NTRs like the Triton.

[kunok]

51 minutes ago, shynung said:

This design is equivalent to two combustion chambers, each fed separately with hydrazine and HTP, feeding a shared nozzle. The idea is that the products of the combustion chamber was to react in the nozzle, producing more energy.

Also, because of the aerospike nozzle, the combustion products essentially react in a nozzle in which half of it is comprised of the ambient atmosphere, the other half being the spike itself. This could work as a decent nozzle/chamber contraption, but its effectiveness decreases as ambient pressure falls, which means the secondary combustion's energy is captured less by the nozzle itself.

I agree with that, as I was agreeing in the previous one, but I think you overestimate the efficiency, I think the pressure will be bad enough. Think about this, IIRC with higher chamber pressure the aerospike works worse. But I'm not sure, maybe the efficiency isn't that bad.

34 minutes ago, sevenperforce said:

As I understand the physics/chemistry, this is going to be a case of shockwave-induced supersonic ram combustion, which is almost impossibly difficult to achieve in scramjets (which must mix a supersonic airflow with a subsonic fuel flow), but will be comparatively straightforward here because both exhaust flows are supersonic. The difference in speed will result in a standing shockwave where they meet; careful shaping of this standing wave by adjustment of the thruster expansion ratios and alignment can be used to ensure complete combustion.

The shape of the aerospike might need to be slightly suboptimal in order to ensure the proper thrust vector, but this will only bleed a few percent of vacuum ISP at most. I'm about 98% sure that an aerospike surface can be used as a combustion chamber in supersonic flow conditions.

The flow pressure of the propellants serves to confine and direct the combustion. The monopropellants have different molar masses and speeds, meaning that their momentum can be configured to ensure proper combustion as well.

Uhh, so to do this, we need a good supersonic combustion of something that it isn't Hydrogen, a good understanding of the shock waves produced in this design, a design that needs to have a stable shock wave front in a huge ratio of external pressures, no turbulence and so on.

A monopropellant engine is a low pressure one, it's a pressure fed design usually, that's the reason they are light. I don't think you can direct and confine the combustion with them.

Are you really taking into account than a lot of the exhaust of the monopropelant engines isn't useful as fuel in the next combustion?

[sevenperforce]

47 minutes ago, kunok said:

Uhh, so to do this, we need a good supersonic combustion of something that it isn't Hydrogen, a good understanding of the shock waves produced in this design, a design that needs to have a stable shock wave front in a huge ratio of external pressures, no turbulence and so on.

A monopropellant engine is a low pressure one, it's a pressure fed design usually, that's the reason they are light. I don't think you can direct and confine the combustion with them.

Are you really taking into account than a lot of the exhaust of the monopropelant engines isn't useful as fuel in the next combustion?

Supersonic combustion is a challenge, yes, but it's not like trying to design a scramjet. With a scramjet, you have very little control over most of what's going into the whole equation. With this design, you have total control over virtually all the variables.

There's an important difference between chamber pressure and flow pressure. Chamber pressure of a monopropellant thruster is typically pretty low, simply because of the application it's used for, though this can be improved with full-flow precombustion turbopumping. Flow pressure, on the other hand, is going to be high because the higher density of the propellant results in a high fluid momentum.

And yes, I'm definitely taking into account the fuel/propellant fractions, as you'll see from the maths that follow.

Anyway, as promised: maths.

I decided to go ahead and calculate the maximum possible vacuum Isp and then dial back from there; getting a feel for maximum possible performance is better than trying to be conservative and adjusting wildly as you go.

These Wikipedia tables helpfully gave a vacuum exhaust velocity of 1.86 km/s for 100% hydrogen peroxide and a dizzying 4.462 km/s for H₂/LOX at a 4.86 mass ratio. Astronautix cites a peak 239 s vacuum specific impulse for certain hydrazine thrusters. Of course these are higher than what would probably be achievable, but again, I'm trying to get a peak characteristic specific impulse for the fuel/combustion system.

I wasn't able to find an exhaustive (pun almost intended) table relating exhaust velocity to mass ratio for H₂/LOX, but picking the optimized 4.83 mass ratio from Wikipedia is probably going to provide the best performance in the end, so that's what I decided to go with. Given the molar masses of hydrogen and oxygen, an H₂:LOX ratio of 1:4.83 can be achieved at a hydrazine:peroxide molar ratio of 1:1.21.

One mole of hydrazine has a mass of 32 grams; its vacuum exhaust velocity of 2.345 km/s means it has a kinetic energy of 88 kJ/mol. One mole of peroxide has a mass of 34 grams; its vacuum exhaust velocity of 1.86 km/s yields a kinetic energy of 71 kJ for 1.21 moles. The total kinetic energy of the combined flow at this ratio will be 159 kJ per mole of hydrazine, corresponding to a flow velocity of 2.09 km/s. So far, so good.

At this ratio, the fluid flow masses 73 grams per mole of hydrazine and contains 23.3 grams of H₂/LOX at the optimal ratio. Combustion of those 23.3 grams at the vacuum exhaust velocity of 4.462 km/s adds 232 kJ to the fluid flow. Distributed within that 73 grams, this adds 2.521 km/s to the flow velocity, resulting in a net flow speed of 4.607 km/s and a specific impulse of 469.6 seconds.

Of course, this is all peak performance. Using solely SL performance and capping the initial flow velocity at the flow speed of peroxide exhaust (e.g., if the excess velocity of the hydrazine is wasted entirely), the same math produces a lower-bound specific impulse of 379 seconds. But that's still not bad for high-density, room-temperature-liquid propellants. This design could very easily be adapted for air augmentation to compensate for this as well, and since the vacuum exhaust velocity is just 3.2 km/s short of orbital velocity, air augmentation will be possible at far higher speeds than would be realistic for conventional rockets. A spaceship with this rocket engine running at optimal specific impulse would only need a remaining propellant mass fraction of 44.8% for the vacuum orbital insertion burn. 

Edited February 16, 2016 by sevenperforce
correction

[sevenperforce]

I haven't been able to find much in the way of theoretical maximum-efficiency thrust-to-weight ratios for monopropellant thrusters, but if it they are anywhere in the 100:1 range (median for medium-efficiency bipropellant engines), then with the supersonic combustion stage of the burn increasing the flow momentum (and, correspondingly, the thrust) by 120%, we'd be looking at a sum T/W ratio of 220:1, which is...dizzying.

At the proper ratio, this fuel combination has a density of 1.255 g/cc. If the Space Shuttle's payload bay were replaced with internal fuel tanks, the total volume would be about 565 cubic meters, containing 710 tonnes of fuel. Sans SSMEs, the Shuttle massed just 59 tonnes. With a 7-tonne aerospike engine of the dual-monocombustion design, the Shuttle Orbiter could launch on its own with a vehicle T/W ratio of 2 and an onboard dV of 11.3 km/s.

[Elthy]

This idea sounds like its breaking the conservation of energy, im just not sure at which point. Propably this:

4 hours ago, Steel said:

• Just putting two streams of bi-propellant together and igniting them you get a reaction where the distribution of momentum is essentially random (i.e the reaction exhausts in all directions). This means there will be some exhaust travelling at 4.3 km s^-1, but equally there will be some travelling back against the flow at 0.7 km s^-1 ,thus disrupting the flow, and some traveling out sideways from the aerospike. If this is the case then you'll not gain any advantage.

 

[K^2]

3 minutes ago, Elthy said:

This idea sounds like its breaking the conservation of energy

It absolutely does. Energy out is greater than energy in. The reason it breaks down is because you should be adding energies, not velocities. This will bring it down to something in the 300s range.

Edited February 16, 2016 by K^2

[sevenperforce]

50 minutes ago, K^2 said:

It absolutely does. Energy out is greater than energy in. The reason it breaks down is because you should be adding energies, not velocities. This will bring it down to something in the 300s range.

Well, dammit. How did I miss that?

Vacuum exhaust velocity ends up being 3,273 m/s, for a specific impulse of 334 seconds. Not nearly so fantastic as formerly expected, but still quite viable.

EDIT:

Actually, on second thought, I'm not so sure.

Chemical energy is not the limiting factor here; chemical rocket propellants have an insane amount of potential energy. After all, only a small amount of the potential energy of a chemical rocket's fuel actually ends up as the kinetic energy of the rocket itself. The limiting factor is the conversion of thermal energy into kinetic energy, and dynamic pressure plays a huge role in that. I don't think conservation of energy is the operative principle here, even if that seems like the easiest way to crunch the numbers.

Kind of how the Oberth effect seems to violate conservation of energy at first glance, but in reality it doesn't.

Edited February 17, 2016 by sevenperforce
changed mind

[DerekL1963]

3 hours ago, sevenperforce said:

The key here is that the standing wave between the two supersonic flows re-creates a new density/velocity maximum which exists within an already-moving reference frame, allowing exhaust velocity to stack. At least, that's how I've designed it; I have no way of knowing whether it will actually work.

Not knowing your qualifications, so no offense intended;   If you can write it up and post it somewhere that isn't the forums (a Google doc?  a blog post?), I'd love to point some real rocket scientists that I know towards it.

[K^2]

2 hours ago, sevenperforce said:

Chemical energy is not the limiting factor here

I_SP is primarily limited by chemical energy. There can be other limiting factors, but maximum vacuum I_SP can be close to theoretical maximum. Things like aerospike are an example of ways to get around some of these limiting factors and get closer to maximum. But if your I_SP comes out higher than thermodynamic maximum, something went wrong. You can bet on it.

[mikegarrison]

I feel compelled to point out that the entire idea of "dual monopropellants" is an oxymoron. Regardless of where the chemical reaction takes place, you have two chemicals in your rocket so neither one is a "monopropellant".

Also, as a point of example, to look at an example of how things work when you burn propellant in a nozzle rather than a combustion chamber you need look no further than an augmentor (aka. afterburner).

[DerekL1963]

2 hours ago, mikegarrison said:

Also, as a point of example, to look at an example of how things work when you burn propellant in a nozzle rather than a combustion chamber you need look no further than an augmentor (aka. afterburner).

Well, no.   The working principle is entirely different in that most jet engines don't actually have a nozzle anything like a rockets.  (And for those that do, it's right at the exit of the tailpipe - downstream of the afterburner.)

Edited February 17, 2016 by DerekL1963

[mikegarrison]

6 minutes ago, DerekL1963 said:

Well, no.   The working principle is entirely different in that most jet engines don't actually have a nozzle.  (And for those that do, it's right at the exit of the tailpipe - downstream of the afterburner.)

I'm sorry, but what are you talking about?

(Look, I know that argument from authority doesn't mean someone is correct, but before we go further I'll mention that I used to work for GE aircraft engines in the combustor and augmentor group. FYI.)

Jet engines definitely have a nozzle. The flow is expanded out to (as much as possible) match the static pressure of the jet to the static pressure of the ambient air. That's what a nozzle is! And augmentors inject fuel into that nozzle, after all the turbines. So....

Edited February 17, 2016 by mikegarrison

[DerekL1963]

2 hours ago, mikegarrison said:

I'm sorry, but what are you talking about?

(Look, I know that argument from authority doesn't mean someone is correct, but before we go further I'll mention that I used to work for GE aircraft engines in the combustor and augmentor group. FYI.)

Jet engines definitely have a nozzle. The flow is expanded out to (as much as possible) match the static pressure of the jet to the static pressure of the ambient air. That's what a nozzle is! And augmentors inject fuel into that nozzle, after all the turbines. So....

OK, I didn't know that.   Though it seems to me that either way, the effective nozzle of the afterburner section would be the exit plane of the tailpipe.

Either way, rocket designers do their damnedest to ensure combustion is complete before the throat for maximum efficiency.  (So sayeth a friend who designs and builds rockets for a living.)

[mikegarrison]

15 minutes ago, DerekL1963 said:

OK, I didn't know that.   Though it seems to me that either way, the effective nozzle of the afterburner section would be the exit plane of the tailpipe.

Either way, rocket designers do their damnedest to ensure combustion is complete before the throat for maximum efficiency.  (So sayeth a friend who designs and builds rockets for a living.)

On a typical gas turbine, the most restrictive throat is the first stage turbine nozzle. And the combustion should be all done in the combustor, before then. As the gas goes through the turbine, it is expanded and work is extracted from it. Then (assuming it's a propulsion jet engine) it goes through a nozzle to further expand it and match the static pressure of ambient. Now since the static pressure of ambient can be wildly different in different flight regimes, there are obviously some compromises to be made. Either that, or you use variable area nozzles. (Turbo fans also have nozzles for the fan air, but there is no combustion going on there.) And if you look at the typical jetliner engine, you just might see a nozzle plug that reminds you a while lot of an aerospike....

Now when you have an afterburner, it has a big ol' second combustion chamber area that is aft of the turbines. Depending on stuff (variable nozzles, the flight regime, etc.) the nozzle after the afterburner area might be convergent or it might be divergent.

But anyway, for the purpose of the discussion in this thread, which was about what are the thermodynamics of having a second combustion happen after you have already had a first combustion earlier, I would think that looking at afterburners and how they work would be useful. Essentially, they end up working in a thermodynamic cycle that is a lot more like a ramjet than a gas turbine (which kind of makes sense because there is no turbine behind them).

 

Oh, and ps. afterburners are not generally efficient! They are used for pure thrust augmentation. However, there are certain cycles where they actually can be efficient. The SR-71 has probably the most complicated engine cycle ever devised. It basically can shift from a pure afterburning turbojet to an almost total ramjet. It's worth learning about.

Edited February 17, 2016 by mikegarrison