Dual monopropellant supersonic combustion rocket

[DerekL1963]

49 minutes ago, sevenperforce said:

The nozzle would likely be actively cooled, probably by the hydrogen peroxide due to its terrific heat capacity.

H2O2 disassociates (very energetically) at around 300 °F (150  C°).   Your cooling system will have to be very carefully designed, even if dealing with yet another thing that can cause H2O2 to disassociate to your list of woes is a good idea in the first place. 

[sevenperforce]

1 hour ago, DerekL1963 said:

H2O2 disassociates (very energetically) at around 300 °F (150  C°).   Your cooling system will have to be very carefully designed, even if dealing with yet another thing that can cause H2O2 to disassociate to your list of woes is a good idea in the first place. 

Active cooling could be used as a preignition source to self-pump the peroxide into its combustion chamber. 

[DerekL1963]

6 minutes ago, sevenperforce said:

1 hour ago, DerekL1963 said:

H2O2 disassociates (very energetically) at around 300 °F (150  C°).   Your cooling system will have to be very carefully designed, even if dealing with yet another thing that can cause H2O2 to disassociate to your list of woes is a good idea in the first place. 

Active cooling could be used as a preignition source to self-pump the peroxide into its combustion chamber. 

o.0   It doesn't actually work that way.   On top of which, you'll need to make your nozzle and chamber cooling passages much more structurally robust to withstand the increased temperatures and pressures post disassociation.  (As well as somehow rigging things so the disassociation doesn't propagate upstream.)  And you'll need cat packs anyhow to start the engine and bring it up to the temperature where it auto-dissociates in the cooling passages.  (Which auto-dissociation also severely limits your ability to throttle and introduced considerable startup and shutdown lag.)

Easier to just not let the auto-dissociation get a foot in the door in the first place.

[sevenperforce]

10 minutes ago, DerekL1963 said:

o.0   It doesn't actually work that way.   On top of which, you'll need to make your nozzle and chamber cooling passages much more structurally robust to withstand the increased temperatures and pressures post disassociation.  (As well as somehow rigging things so the disassociation doesn't propagate upstream.)  And you'll need cat packs anyhow to start the engine and bring it up to the temperature where it auto-dissociates in the cooling passages.  (Which auto-dissociation also severely limits your ability to throttle and introduced considerable startup and shutdown lag.)

Easier to just not let the auto-dissociation get a foot in the door in the first place.

Yeah, probably right.

[A Fuzzy Velociraptor]

55 minutes ago, sevenperforce said:

Active cooling could be used as a preignition source to self-pump the peroxide into its combustion chamber. 

That would be highly unadvisable. Use of a normal catalyst either platinum or a permangenate once in the chamber would be much more advised.

I will address the rest of the thread since I was on last night when I have a chance to make a long post.

[sevenperforce]

21 hours ago, A Fuzzy Velociraptor said:

I will address the rest of the thread since I was on last night when I have a chance to make a long post.

Definitely interested.

Maybe next week I'll mock up a way to use the sort of design for air-augmentation...

[A Fuzzy Velociraptor]

Anyway, sorry I haven't responded to this in like a really long time. I've been meaning to and have been extremely busy and distracted with my own work. So I think what have a couple holes in your understanding of how rockets work and mistaking some causes. We've already talked about the energy and thermodynamic misunderstandings associated with this design so I'll leave that out.

I think you are misunderstanding the reason monopropellant engines tend to have a high thrust to weight ratio. While they generally don't produce large thrust, they are very light. They can be this light because in general monopropellant engines are very simple. Often times they will be pressure fed and of course only require a single inlet. Often times the same can be said for the small bi-prop engines (though the monoprop is often still lighter).If you compare the MR-104A/C 440N monopropellant thruster and the Hi-Pat 440N High Altitude thuster both made by Aerojet Rocketdyne, the difference in mass is that the MR-104A/C is about about 60% lighter than the Hi-Pat.  Monoprop engines don't release nearly the chemical energy most bi-propellant reactions used in rocket propulsion. As such they may not require the cooling other systems may need, but this also means they don't have nearly the energy the other systems do. If you compare the nozzles of the MR-104A/C and the Hi-Pat, the Hi-Pat has a significant expansion added to it which about doubles the nozzle length and before the expansion the two engines are about the same size. I don't know how much that expansion weighs, but by dropping it the mass of the engine would likely be significantly reduced bringing the two mass figures significantly closer together. A monoprop engine on its own scaled up would likely be slightly lighter than a biprop engine of the same thrust but at significantly lower performance.

Also consider that, while thrust to weight is an important factor, the variation of the thrust to weight of just the engine masses among almost any chemical system is going to be almost irrelevant. For example consider that you desire an initial acceleration of somewhere in the range of 1.25G and both engines listed earlier produce 100lbf of thrust. So vehicle mass will likely come in around to about 80lbm to start. The MR-104A/C weighs 4.11lbm and the Hi-Pat 11.5lbm which is about a 9% difference in the total mass ratio compared to the 60% in the engines. This is a fairly extreme example, the R-4D is a 4% difference. The point here being that seemingly large differences in thrust to weight ratios of engines aren't nearly as large an effect on the whole vehicle as may appear. We do talk about thrust to weight ratios and why they are important but the main reason we talk about them especially with engines is when talking about mass limited vs energy limited propulsion systems. Note the difference in thrust to weight of BPT-200 Hall Effect Thruster vs the chemical example of .0024 and 24.33 which is a difference by about a factor of 10000 and that only factors in the engine and not the masses of all the systems needed to make the engine work which will only hit the energy limited system harder.

While you talk about thrust to weight quite a lot and keeping the mass of the system low, the engine you are proposing is actually quite complex. The reason monoprop engines are so light is largely because of that simplicity. Aerospike/single expansion nozzles are not light apparatus to start with and the air intakes associated with an air augmentation system are significantly more massive. In addition by using two separate combustion chambers you are essentially doubling the weight of the combustion chambers for a given mass flow rate which significantly more hampers your thrust to weight. 

The last thing is in regards to the air augmentation system. Air augmentation is a very complex problem of aerodynamics that honestly I'm nowhere the level of education required to design of fully comment on one. I can tell you that they are very heavy and difficult to design, they can allow for significantly higher specific impulse but are expensive, complex, difficult to design, and will only work in the lower parts of the atmosphere as I understand them. The weight of the system is large enough that it will significantly affect your system rather than the weight of monoprop vs biprop engines. Also given the time and expense of manufacture it would likely be unadvised for a disposable vehicle and would likely have limited advantage is a vertical take off vehicle. Also normally you increase the working mass after combusting the propellant rather than before. Nitrogen especially in certain reactions can act as almost an anti-catalyst and can significantly slow down your reaction. This is already non optimal in a combustion chamber where the fluid may be moving at 50m/s, it is significantly worse when flying down a nozzle moving at 500+m/s. You will likely wind up with some nitrous oxide as well from oxygen colliding with the nitrogen though that shouldn't have too large of an effect except a bit of pollution. At least with your crude drawing the air would likely significantly impede the combustion of your propellants.

Edited April 10, 2016 by A Fuzzy Velociraptor

[todofwar]

Nat sure about the physics of this but from a chemistry standpoint I would rather stand in a room full of hydrogen and oxygen dewars than one bottle of anhydrous hydrazine or peroxide. Those things are not happy being themselves and will happily detonate, hence why hydrazine is a useful rocket fuel. Peroxide is so bad you can't even buy higher than 30% cause it starts spontaneously detonating above that. So if the argument is around storage vs LH2/LO2 than I would say the safety concerns would outweigh other benefits.

[Steel]

11 minutes ago, todofwar said:

Nat sure about the physics of this but from a chemistry standpoint I would rather stand in a room full of hydrogen and oxygen dewars than one bottle of anhydrous hydrazine or peroxide. Those things are not happy being themselves and will happily detonate, hence why hydrazine is a useful rocket fuel. Peroxide is so bad you can't even buy higher than 30% cause it starts spontaneously detonating above that. So if the argument is around storage vs LH2/LO2 than I would say the safety concerns would outweigh other benefits.

Peroxide is nowhere near as bad as you make it sound, if stored correctly its pretty harmless. There are companies on the internet that will sell qualified buyers 80-90% "Propellant Grade Peroxide" in jerry cans loaded onto pallets.

Also hydrazine is not used as a rocket fuel due to its propensity to detonate when just sat around. It may be toxic but it does not just explode at random intervals (especially when correctly stored), otherwise you would never even think about using it as a storable propellant in a missile (as it was all through the cold war).

Edited April 10, 2016 by Steel

[todofwar]

24 minutes ago, Steel said:

Peroxide is nowhere near as bad as you make it sound, if stored correctly its pretty harmless. There are companies on the internet that will sell qualified buyers 80-90% "Propellant Grade Peroxide" in jerry cans loaded onto pallets. Also hydrazine is not used as a rocket fuel due to its propensity to detonate when just sat around because it may be toxic but it does not just explode when correctly stored, otherwise you would never even think about using it as a storable propellant in a missile (as it was all through the cold war).

Honestly had no idea they used peroxide as fuel. I knew about hydrazine, but since we have anhydrous hydrazine in my lab I'm well aware its fine when stored properly, we always treat it like a bomb mostly because we all know it's used as rocket fuel. But peroxide is something that is too dangerous to be used as a reagent. If we're talking about benefits vs H2/O2, I would say since that's your overall reaction anyway might as well stick to that cause I doubt the storage of that much peroxide will be any safer or easier. 

And I'm going to modify my previous statement from a room full of dewars to a room full of hydrogen and oxygen tanks, just remembered the whole boiling some off to keep themselves cold thing and that would be bad to have them in one room. 

Edited April 10, 2016 by todofwar

[Kryten]

6 minutes ago, todofwar said:

Honestly had no idea they used peroxide as fuel.

They generally don't, though it has been done. The British BLACK ARROW rocket from the 60's used kerosene/peroxide for the first two stages, but only recent use has been a kerolox/peroxide third stage and peroxide monoprop roll thrusters on the new Chinese CZ-6 light rocket.

[Steel]

9 minutes ago, todofwar said:

Honestly had no idea they used peroxide as fuel.

Yeah if you're interested you can read about the Black Arrow rocket, which was a British rocket with engines burning high test peroxide (i.e hihgly concentrated peroxide) and RP-1.

[A Fuzzy Velociraptor]

10 hours ago, todofwar said:

Nat sure about the physics of this but from a chemistry standpoint I would rather stand in a room full of hydrogen and oxygen dewars than one bottle of anhydrous hydrazine or peroxide. Those things are not happy being themselves and will happily detonate, hence why hydrazine is a useful rocket fuel. Peroxide is so bad you can't even buy higher than 30% cause it starts spontaneously detonating above that. So if the argument is around storage vs LH2/LO2 than I would say the safety concerns would outweigh other benefits.

The first part isn't really a logical argument especially when talking about rocket propellants. These things have to have a lot a chemical energy which can be released from a reaction, they aren't necessarily going to be extremely human friendly. Also as the other fellow said, peroxide is actually quite stable at high concentrations and actually gets more stable on its own as the concentrations rise. Yes HTP can start decomposing which causes a run away reaction rather than a detonation. HTP itself is actually decently stable and requires a significant amount of energy to begin to thermally decompose. Of course is certain impurities get into the system the impurity can act as a catalyst which can have non-optimal consequences. HTP is often used in some capacity as a propellant though often in monopropellant and normally not main engines for space launch vehicles. It can be bought in 50, 70, 90, 98 and higher concentrations from various suppliers though you may have to buy a large quantity of the stuff depending on the supplier.

Also the safety concerns of HTP and Hydrazine can outweigh the storage concerns of hydrolox. For one safety concerns can easily be managed and mitigated where storability may not. There are a number of other reasons and variations why some entities pursue hydrolox vs kerolox vs keroperox vs UDMH/NTO etc and it can't really be simplified down to a couple sentence discussion.

[todofwar]

3 minutes ago, A Fuzzy Velociraptor said:

The first part isn't really a logical argument especially when talking about rocket propellants. These things have to have a lot a chemical energy which can be released from a reaction, they aren't necessarily going to be extremely human friendly. Also as the other fellow said, peroxide is actually quite stable at high concentrations and actually gets more stable on its own as the concentrations rise. Yes HTP can start decomposing which causes a run away reaction rather than a detonation. HTP itself is actually decently stable and requires a significant amount of energy to begin to thermally decompose. Of course is certain impurities get into the system the impurity can act as a catalyst which can have non-optimal consequences. HTP is often used in some capacity as a propellant though often in monopropellant and normally not main engines for space launch vehicles. It can be bought in 50, 70, 90, 98 and higher concentrations from various suppliers though you may have to buy a large quantity of the stuff depending on the supplier.

Also the safety concerns of HTP and Hydrazine can outweigh the storage concerns of hydrolox. For one safety concerns can easily be managed and mitigated where storability may not. There are a number of other reasons and variations why some entities pursue hydrolox vs kerolox vs keroperox vs UDMH/NTO etc and it can't really be simplified down to a couple sentence discussion.

I'm willing to grant that since my experience with it is in a synthetic chemistry setting where we will need to dip needles and other things into the bottle to get the reagent out, I may have a warped perception of its stability. No one I know touches anything higher than 30% peroxide because they worry about it detonating, while anhydrous hydrazine just gets a little extra caution. 

[sevenperforce]

15 hours ago, todofwar said:

I'm willing to grant that since my experience with it is in a synthetic chemistry setting where we will need to dip needles and other things into the bottle to get the reagent out, I may have a warped perception of its stability. No one I know touches anything higher than 30% peroxide because they worry about it detonating, while anhydrous hydrazine just gets a little extra caution. 

As others have said, aerospace-grade propellants are already usually pretty nasty stuff, so this comes with the territory. Granted, there are certain things which exceed even rocket scientists' love for delta-v, like pentaborane, but neither HTP nor hydrazine are nearly that bad. Hydrazine is fairly stable in the absence of a catalyst or ignition source; it's just quite toxic. HTP is stable enough if stored correctly and you won't be dipping anything into the tank.

On 4/10/2016 at 10:09 PM, A Fuzzy Velociraptor said:

Anyway, sorry I haven't responded to this in like a really long time. I've been meaning to and have been extremely busy and distracted with my own work. So I think what have a couple holes in your understanding of how rockets work and mistaking some causes. We've already talked about the energy and thermodynamic misunderstandings associated with this design so I'll leave that out.

I think you are misunderstanding the reason monopropellant engines tend to have a high thrust to weight ratio. While they generally don't produce large thrust, they are very light. They can be this light because in general monopropellant engines are very simple. Often times they will be pressure fed and of course only require a single inlet. A monoprop engine on its own scaled up would likely be slightly lighter than a biprop engine of the same thrust but at significantly lower performance.

While you talk about thrust to weight quite a lot and keeping the mass of the system low, the engine you are proposing is actually quite complex. The reason monoprop engines are so light is largely because of that simplicity. Aerospike/single expansion nozzles are not light apparatus to start with and the air intakes associated with an air augmentation system are significantly more massive. In addition by using two separate combustion chambers you are essentially doubling the weight of the combustion chambers for a given mass flow rate which significantly more hampers your thrust to weight. 

The last thing is in regards to the air augmentation system. Air augmentation is a very complex problem of aerodynamics that honestly I'm nowhere the level of education required to design of fully comment on one. I can tell you that they are very heavy and difficult to design, they can allow for significantly higher specific impulse but are expensive, complex, difficult to design, and will only work in the lower parts of the atmosphere as I understand them. The weight of the system is large enough that it will significantly affect your system rather than the weight of monoprop vs biprop engines.

Yeah, the whole concept (at least as provided in the OP) was pretty much dead in the water as soon as I realized that my energy calculations were totally hacked due to improperly stacking exhaust velocities.

The simplest air-augmentation setup is to wrap a simple shroud around the exhaust nozzle, which can increase thrust by up to 15% at launch due to pure ejector effect and will ramp up considerably at higher velocities. Using a single or double monopropellant injection a la TAN (thrust-augmented nozzle) arrangement could give the launch thrust boost desired while allowing a wide range of fuel-air-oxidizer mixture ratios throughout the flight path.

[Ravenchant]

There have been several mentions of "monopropellant engines having high TWRs." Any examples? From what I know their thrust to weight ratios are abysmal compared to bipropellant engines, and the highest I could find is Aerojet Rocketdyne's MR-80 thruster at a whopping 37.1.

[A Fuzzy Velociraptor]

On 4/11/2016 at 3:01 PM, Ravenchant said:

There have been several mentions of "monopropellant engines having high TWRs." Any examples? From what I know their thrust to weight ratios are abysmal compared to bipropellant engines, and the highest I could find is Aerojet Rocketdyne's MR-80 thruster at a whopping 37.1.

I believe seven was referring to engines of similar thrust levels. If you take a look through Aerojet Rocketdyne's capabilities page most of the monopropellant engines are lighter than the bipropellant engines of similar thrust levels. Monopropellant engines tend not to be used for high thrust applications so the thrust to weight scaling is a little screwy.