Dual monopropellant supersonic combustion rocket

[11of10]

Considering that Hydrazine (N2H4) used in monopropellant thrusters breaks down into Ammonia (NH3) Hydrogen (H2) and Nitrogen (N2) and that Hydrogen Peroxide used in monopropellant thrusters breaks down into H2O and some O2, you're essentially getting a H2 + O2 reaction, possibly without stochiometric ratios, and outside of a combustion chamber, which essentially makes it a very expensive and inefficient, not to mention toxic lighter. Or a blowtorch.

 

http://www.mtc.edu.eg/ASAT13/pdf/PP22.pdf <- Hydrazine monopropellant thruster

http://www.esa.int/gsp/ACT/doc/PRO/ACT-RPR-PRO-JPC2006-HP%20Rockets%202006-5239.pdf <- Hydrogen peroxide monopropellant thruster

[B787_300]

I would like to point out that the burning of fuels when not in a combustion chamber is non optimal.  The reason rockets have combustion chambers is so that all the reactants can react.  They have to be a certain length based off the chemical reactions L* value and increasing the temperature and pressure in the combustion chamber increase the rate of reaction.  If you are trying to burn them outside it will be much harder to control.  Also you will have a LOT of unburned propellants in the design that you are proposing.  and Hydrazine is not something you want to let escape unburned.

[sevenperforce]

I wonder if dinitrogen pentoxide salt could be dissolved into high-test hydrogen peroxide to increase both its monopropellant impulse and its oxidizer ratio.

16 hours ago, DerekL1963 said:

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.

Thanks, I'll try to do that. As far as qualifications -- I've got a research degree in the hard sciences, but rocket engineering is just a hobby.

14 hours ago, K^2 said:

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.

ISP ultimately depends on chemical energy, yes, but it is much more directly sensitive to variables like chamber pressure. The larger your pressure drop, the better your ISP. I don't think the ~470 seconds is anywhere close to the thermodynamic maximum for these fuels; a tripropellant fluorine-LH₂-lithium rocket can get upwards of 540 seconds. Injecting a very small amount of liquid hydrogen into a kerolox engine can boost the specific impulse from 330 seconds to 415 seconds because of how it changes the expansion rate and ratio of the propellant, not because the hydrogen adds significantly more chemical potential energy.

The peak pressure of a subsonic choked flow is distributed in every direction and places tremendous stress on your combustion chamber; the peak pressure of a supersonic flow is limited to the flow axis and results in vastly-reduced stress loading. I see no reason why a second combustion cannot achieve extremely high efficiencies via uniaxial supersonic compression and stack velocities.

14 hours ago, mikegarrison said:

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).

I used the term "dual monopropellant combustion" because the two fuels are accelerated through monopropellant thrusters before being combusted together at supersonic speeds. If you can think of a better descriptor, I'd be interested!

An afterburner suffers from low efficiency due to the inherently poor compression ratio of combining a supersonic gas flow (usually of oxidizer) with a subsonic aerosolized liquid flow (usually fuel). When both fuels are highly-tuned supersonic gas flows, stupidly high compression ratios can be achieved.

2 hours ago, 11of10 said:

Considering that Hydrazine (N2H4) used in monopropellant thrusters breaks down into Ammonia (NH3) Hydrogen (H2) and Nitrogen (N2) and that Hydrogen Peroxide used in monopropellant thrusters breaks down into H2O and some O2, you're essentially getting a H2 + O2 reaction, possibly without stochiometric ratios, and outside of a combustion chamber, which essentially makes it a very expensive and inefficient, not to mention toxic lighter. Or a blowtorch.

Hydrazine will break down into pure diatomic nitrogen and diatomic oxygen when the chamber temperature is raised high enough. So although the fuel is toxic, the products are not.

The design uses a supersonic combustor "chamber" for secondary combustion; one side is formed by the aerospike nozzle while the other side of the "chamber" is formed by the standing shockwave where the two supersonic propellant flows meet. Thus, a high compression ratio can be achieved.

2 hours ago, B787_300 said:

Also you will have a LOT of unburned propellants in the design that you are proposing.  and Hydrazine is not something you want to let escape unburned.

The hydrazine will flash completely to nitrogen and oxygen (and maybe a little ammonia) in its monoprop thrust chamber; the hydrogen and oxygen from H₂O₂ decomposition will burn quite merrily. At most you'll have some leftover unburned hydrogen in the exhaust.

Picking a lower hydrazine combustion temperature, resulting in some ammonia in the flow, might actually be better. At the right ratios, you could get ammonia, oxygen, and hydrogen in and nothing but water and diatomic nitrogen out, causing a huge energy boost from the highly-energetic nitrogen-nitrogen reaction.

[fredinno]

3 hours ago, B787_300 said:

I would like to point out that the burning of fuels when not in a combustion chamber is non optimal.  The reason rockets have combustion chambers is so that all the reactants can react.  They have to be a certain length based off the chemical reactions L* value and increasing the temperature and pressure in the combustion chamber increase the rate of reaction.  If you are trying to burn them outside it will be much harder to control.  Also you will have a LOT of unburned propellants in the design that you are proposing.  and Hydrazine is not something you want to let escape unburned.

Could some sort of staged combustion allow for two combustion reactions at once?

[sevenperforce]

12 minutes ago, fredinno said:

Could some sort of staged combustion allow for two combustion reactions at once?

Well, hydrazine and peroxide are merrily hypergolic together, but with a very poor ISP. Combustion against a diverging nozzle via shockwave-induced compression of a supersonic flow is the only way I can think of to have a chance of stacking velocities.

A supersonic flow has the advantage of not being able to have back-propagation.

[A Fuzzy Velociraptor]

While I can't say I know much about supersonic combustion and the instabilities associated with it, I can say there are some serious errors in the way you are estimating your ISP. Velocities do not add linearly here because you are not adding additional velocity, you are adding energy and representing that energy in terms of a velocity.

The exit velocity of a nozzle for a one dimensional analysis assuming 100% efficiency may be represented by 

There are several ways you can increase the effective velocity. Raising the temperature, decreasing the average molar mass of your reaction products,  decreasing the ratio of specific heats, and increasing the pressure in chamber to exit pressure will increase your effective velocity. Decreasing the molar mass of the products normally come with a decrease in gamma since products with fewer atoms per molecule will have a lower gamma values.

The stoichiometric ratio for a hydrazine/HTP reaction is as follows:

N2H4+2(H2O2)<=>4(H2O)+N2

This leads to a O/F ratio of 2.125. Normally engines tend to run slightly fuel rich because it tends to be a bit more efficient and reduces temperature and wear on engines from hot ozidizer (though from experience using CEA H2O2 based propellant combinations tend to prefer to run oxidizer rich).

Now to give you the following numbers I had to assume that your reaction occurs inside the combustion chamber rather than outside it. I expect these numbers will be substantially higher than you would actually experience in your design due to less complete mixing and combustion issues.

Inputs into CEA

prob case=12349701 ro equilibrium 

 ! iac problem
o/f 1.9 1.95 2 2.05 2.1 2.125 2.15 2.2 2.25 2.3
p,bar  100
pip 100 500 5000 50000 5000000
reac
 fuel  N2H4 wt%=100 t,k=298.15
 oxid H2O2 wt%=100 t,k=298.15
output    short
output trace=1e-5
end

I assumed the chamber pressure to be 10MPa. Yours may be higher or lower, likely lower if you intend to pressure feed your system. I assumed you are using pure Hydrazine and HTP rather than a lower grade. I also assumed both would be a room temperature. In reality you would want to use the HTP to cool your nozzle so your temperature will likely be more than 300K though I would suggest not exceeding 400K.

              THEORETICAL ROCKET PERFORMANCE ASSUMING EQUILIBRIUM

           COMPOSITION DURING EXPANSION FROM INFINITE AREA COMBUSTOR

 Pin =  1450.4 PSIA
 CASE = 12349701       

             REACTANT                    WT FRACTION      ENERGY      TEMP
                                          (SEE NOTE)     KJ/KG-MOL      K  
 FUEL        N2H4                         1.0000000     95180.000    298.150
 OXIDANT     H2O2                         1.0000000   -135880.000    298.150

 O/F=    2.12500  %FUEL= 32.000000  R,EQ.RATIO= 0.999511  PHI,EQ.RATIO= 0.999022

                 CHAMBER   THROAT     EXIT     EXIT     EXIT     EXIT     EXIT
 Pinf/P            1.0000   1.7350   100.00   500.00  5000.00  50000.0 5000000.
 P, BAR            100.00   57.637   1.0000  0.20000  0.02000  0.00200  0.00002
 T, K             3235.20  3051.29  1796.75  1345.60   837.48   489.25   215.97
 RHO, KG/CU M    7.1174 0 4.3891 0 1.3387-1 3.5782-2 5.7492-3 9.8412-4 2.3416-5
 H, KJ/KG        -1765.97 -2514.57 -6646.54 -7693.44 -8724.09 -9346.60 -9920.28
 U, KJ/KG        -3170.98 -3827.77 -7393.55 -8252.39 -9071.96 -9549.83 -10005.7
 G, KJ/KG        -44060.6 -42404.9 -30135.9 -25284.8 -19672.7 -15742.7 -12743.7
 S, KJ/(KG)(K)    13.0732  13.0732  13.0732  13.0732  13.0732  13.0732  13.0732

 M, (1/n)          19.145   19.319   19.998   20.016   20.016   20.016   21.024
 MW, MOL WT        19.145   19.319   19.998   20.016   20.016   20.016   20.016
 (dLV/dLP)t      -1.01584 -1.01259 -1.00032 -1.00001 -1.00000 -1.00000 -4.75855
 (dLV/dLT)p        1.3127   1.2638   1.0113   1.0003   1.0000   1.0000 107.9642
 Cp, KJ/(KG)(K)    5.4074   5.0738   2.5273   2.1720   1.8833   1.69851205.4899
 GAMMAs            1.1397   1.1401   1.2018   1.2366   1.2830   1.3237   1.0701
 SON VEL,M/SEC     1265.4   1223.6    947.5    831.4    668.1    518.7    302.3
 MACH NUMBER        0.000    1.000    3.297    4.141    5.584    7.507   13.358

 PERFORMANCE PARAMETERS

 Ae/At                      1.0000   12.841   43.592   250.41  1401.51  56793.1
 CSTAR, M/SEC               1862.0   1862.0   1862.0   1862.0   1862.0   1862.0
 CF                         0.6571   1.6779   1.8491   2.0034   2.0911   2.1688
 Ivac, M/SEC                2296.8   3363.4   3605.4   3823.7   3945.9   4059.5
 Isp, M/SEC                 1223.6   3124.3   3443.1   3730.4   3893.7   4038.4

 MOLE FRACTIONS

 *H              7.198 -3 5.148 -3 1.357 -5 2.447 -8 9.519-16 3.319-30 0.000  0
 HO2             7.026 -5 4.143 -5 4.930 -8 6.141-10 2.574-13 1.928-19 0.000  0
 *H2             5.419 -2 4.480 -2 1.512 -3 2.398 -5 1.099-10 4.616-21 0.000  0
 H2O             6.8792-1 7.0942-1 7.9741-1 7.9988-1 7.9992-1 7.9992-1 7.5201-1
 H2O2            1.977 -5 1.123 -5 1.946 -8 3.919-10 3.772-13 2.482-18 1.171-33
 *NO             7.834 -3 6.014 -3 1.242 -4 8.189 -6 5.696 -8 5.014-12 2.347-24
 NO2             1.014 -5 6.205 -6 1.899 -8 1.029 -9 5.153-11 5.471-13 1.367-18
 *N2             1.8725-1 1.8990-1 1.9964-1 1.9988-1 1.9988-1 1.9988-1 1.9988-1
 *O              2.832 -3 1.960 -3 3.245 -6 1.238 -8 4.304-14 9.658-25 0.000  0
 *OH             3.864 -2 3.045 -2 5.373 -4 1.563 -5 4.598 -9 6.702-16 8.444-37
 *O2             1.4017-2 1.2233-2 7.5703-4 1.9968-4 1.9567-4 1.9570-4 1.9570-4
 H2O(cr)         0.0000 0 0.0000 0 0.0000 0 0.0000 0 0.0000 0 0.0000 0 4.7917-2

  * THERMODYNAMIC PROPERTIES FITTED TO 20000.K

 NOTE. WEIGHT FRACTION OF FUEL IN TOTAL FUELS AND OF OXIDANT IN TOTAL OXIDANTS

I was going to put the rest of the CEA file in a spoiler but it does not appear to keep the formatting. If it is desired I can paste the whole file here but it will be rather long.

With an aerospike nozzle, and most altitude compensating nozzles your expansion ratio will vary such that the exit pressure will be equal to the ambient pressure. Your maximum expansion ratio will be roughly equal to your geometric area ratio (the area of the base of your nozzle divided by the throat area). This ratio likely won't be significantly more than the 250 expansion ratio.

One dimensional analysis doesn't apply very well to your system directly due to various inefficiencies in your design. However, even the one dimensional analysis numbers are substantially below your target numbers. Now your propellant combination isn't a bad one and used in a more conventional chamber design could experience quite low structural mass ratios since your propellants have density greater than water. While there would be environmental and cost concerns for safety systems there is currently funding available thought DARPA TTO for high density propellants where this may fit well.

Edited February 17, 2016 by A Fuzzy Velociraptor
Something got deleted

[sevenperforce]

Thanks for all the analysis!

I'm not sure your starting assumptions are right, though. I'm not proposing a hydrazine-peroxide rocket. I'm proposing separate hydrazine and peroxide monopropellant thrusters which each provide an initial impulse, followed by what is essentially a supersonic-flow afterburner to combust the oxygen in the peroxide exhaust stream with the hydrogen in the hydrazine exhaust stream.

The hydrazine thrusters and peroxide thrusters provide an initial impulse to the rocket prior to supersonic mixing and recombustion. Mixing produces a standing termination shock with supersonic flow; recombustion essentially thrusts against this termination shock to provide the second stage of impulse.

Thus, it isn't really appropriate to compare this design to a simple hydrazine-peroxide reaction in a single combustion chamber. Granted, it is the same amount of chemical potential energy, but the chemical reactions are segregated. The high-thrust, low-impulse catalyzed decomposition reactions take place first, with the lower-thrust, higher-impulse 2H2+O2 reaction taking place in the combined supersonic exhaust stream. This design overcomes the primary limiting variable in combustion chamber design: namely, limited chamber pressure and temperature.

Recombustion inside a strongly supersonic exhaust stream should ideally prevent backward propagation of secondary combustion, causing only positive axial acceleration. Unfortunately, my understanding of supersonic flow dynamics and thermodynamic efficiency is not quite advanced enough to know for sure.

Edited February 17, 2016 by sevenperforce

[A Fuzzy Velociraptor]

I am aware that you aren't proposing a hydrazine/HTP engine. Unfortunately I do not have the analytical tools or motivation available to me to right now to fully analyze your proposed system. Under the most ideal circumstances, the maximum energy you will be able to extract will be equal to the one dimensional analysis numbers. Your system will likely have incomplete mixing due to the way your fuel and oxidizer are flowing into each other. Normally you want to have your fuel stream impinging on your oxidizer stream. Here you may only get mixing and combustion along the boundary layer between your two flows which will reduce the total chemical energy extracted. Breaking the combustion into two separate sections wont actually allow you to increase the maximum energy extracted from the combustion. The ISP found in the one dimensional analysis represents the maximum energy you can expect to extract from the system. In reality you should expect a lower amount of energy extraction and substantially lower for your system.

I think I understand what you are trying to do but I think there is a fundamental error in your design that you aren't accounting for. I don't see how you are addressing the issues regarding chamber pressure and chamber temperature. While your chamber temperature will be lower, your nozzle will still have standard heat problems, and the high heat capacity of hydrogen peroxide means that cooling is not much of an issue anyway. Also I think you may be misunderstanding how chamber pressure works. Your chamber pressure would be still required to be quite high and you wouldn't be able to decrease your needed chamber pressure inside the combustion chamber by having the main combustion occurring outside the chamber itself in the same way that you still need to have high pressure in your pipes and injectors to push the stuff into the chamber.

Edited February 17, 2016 by A Fuzzy Velociraptor

[mikegarrison]

The thing that confuses me is this:

There is a thermodynamic cycle advantage to doing your combustion at high pressure. But in order to capture the benefit of that, you need to fully expand the output flow. (Of course in space you can never fully expand to vacuum, but that's not the issue.) The issue is that your second stage of combustion happens in the nozzle. So where do you expand the flow? You are already in the nozzle. Do you have a second nozzle?

[B787_300]

1 hour ago, mikegarrison said:

The issue is that your second stage of combustion happens in the nozzle. So where do you expand the flow? You are already in the nozzle. Do you have a second nozzle?

YOu are correct.  The flow needs to be expanded but the benefit of an aerospike is that the air is one of the "walls" of the nozzle.  the other is the spike itself.

[mikegarrison]

2 minutes ago, B787_300 said:

YOu are correct.  The flow needs to be expanded but the benefit of an aerospike is that the air is one of the "walls" of the nozzle.  the other is the spike itself.

That doesn't answer my question. It just describes the nozzle that the burning is taking place in. The point is that if you have already expanded your flow before you do this secondary combustion, how do you get the benefits from it?

[A Fuzzy Velociraptor]

51 minutes ago, mikegarrison said:

That doesn't answer my question. It just describes the nozzle that the burning is taking place in. The point is that if you have already expanded your flow before you do this secondary combustion, how do you get the benefits from it?

I think in this scenario the flow is partially expanded before it is ignited. There will be some benefit from igniting the flow and letting it further expand though the amount of energy extracted from that system will be lower than just using a conventional bi-propellant combustion chamber and nozzle.

[sevenperforce]

15 hours ago, A Fuzzy Velociraptor said:

I am aware that you aren't proposing a hydrazine/HTP engine. Unfortunately I do not have the analytical tools or motivation available to me to right now to fully analyze your proposed system. Under the most ideal circumstances, the maximum energy you will be able to extract will be equal to the one dimensional analysis numbers. Your system will likely have incomplete mixing due to the way your fuel and oxidizer are flowing into each other. Normally you want to have your fuel stream impinging on your oxidizer stream. Here you may only get mixing and combustion along the boundary layer between your two flows which will reduce the total chemical energy extracted. Breaking the combustion into two separate sections wont actually allow you to increase the maximum energy extracted from the combustion. The ISP found in the one dimensional analysis represents the maximum energy you can expect to extract from the system.

I think I understand what you are trying to do but I think there is a fundamental error in your design that you aren't accounting for.

Also I think you may be misunderstanding how chamber pressure works. Your chamber pressure would be still required to be quite high and you wouldn't be able to decrease your needed chamber pressure inside the combustion chamber by having the main combustion occurring outside the chamber itself in the same way that you still need to have high pressure in your pipes and injectors to push the stuff into the chamber.

That seems increasingly likely.

I was uncertain about the analysis, so I looked up all the individual bond dissociation energies for the individual reactants to get an idea of the total chemical potential energy...I ended up getting just 5.08 MJ/kg for the 1:1.2 molar ratio I had specified earlier, which, if converted completely to kinetic energy and expanded to infinity, would only result in an exhaust velocity of 3.19 km/s. Rather disappointing. I'm still not sure why the staged supersonic combustion wouldn't do what I want it to do, but I suppose that's just how it is.

Or maybe I'm underestimating the chemical potential energy somewhere along the line. I hope I am...

But even if I'm not, I think the design could possibly still hold promise. To your points: I depicted the propellant flows as parallel for the sake of simplicity when rendering the 3D model, but in reality they would probably be staggered and overlap, so as to mix completely.

And as far as the pressure issue is concerned -- the goal would be to use dynamic flow pressure against the nozzle surface in place of hydrostatic pressure inside the combustion chamber. Of course, dynamic flow pressure won't typically be as high as the chamber pressure it came from; high-efficiency bipropellant liquid-fueled engines typically have a chamber pressure anywhere from 20-100 times greater than their dynamic flow pressure. But propellants with higher densities like hydrazine and peroxide will have dynamic pressures closer to their chamber pressure. That doesn't mean the chamber pressure will be negligible; the thrusters would still need turbopumps. But the thrust-to-weight ratio would be greatly improved.

Another consideration which might prove more promising...

If two supersonic flows are being combined and combusted outside of the chamber, then the opportunity for air augmentation becomes extraordinarily good. Consider a setup like this:

The oxidizer flow could be used to compress the flow of atmospheric air and combine it with the exhaust to significantly improve specific impulse...and with a very high thrust-to-weight ratio, this could become viable.

 

[sevenperforce]

17 hours ago, mikegarrison said:

The thing that confuses me is this:

There is a thermodynamic cycle advantage to doing your combustion at high pressure. But in order to capture the benefit of that, you need to fully expand the output flow. (Of course in space you can never fully expand to vacuum, but that's not the issue.) The issue is that your second stage of combustion happens in the nozzle. So where do you expand the flow? You are already in the nozzle. Do you have a second nozzle?

Compare the shape of a ramjet nozzle:

to a scramjet nozzle:

Because the flow is not choked and combustion takes place at a supersonic speed, there is no forward propagation of shockwaves and so the shape is completely different. Although my design isn't technically a scramjet, it's closer to a scramjet than it is to a conventional rocket. Combustion can take place within the expansion nozzle while maintaining a high compression ratio.

[mikegarrison]

34 minutes ago, sevenperforce said:

Compare the shape of a ramjet nozzle:

to a scramjet nozzle:

Because the flow is not choked and combustion takes place at a supersonic speed, there is no forward propagation of shockwaves and so the shape is completely different. Although my design isn't technically a scramjet, it's closer to a scramjet than it is to a conventional rocket. Combustion can take place within the expansion nozzle while maintaining a high compression ratio.

I'm not questioning that the combustion can happen there. I'm questioning how you can get any benefit from it.

[sevenperforce]

5 minutes ago, mikegarrison said:

I'm not questioning that the combustion can happen there. I'm questioning how you can get any benefit from it.

Err...by expansion against the remaining portion of the nozzle?

[mikegarrison]

Just now, sevenperforce said:

Err...by expansion against the remaining portion of the nozzle?

Well that's the thing -- how much of the nozzle is remaining to expand against? That flow is moving fast, and chemistry takes time. How do you make it happen before there is no more nozzle?

[sevenperforce]

Just now, mikegarrison said:

Well that's the thing -- how much of the nozzle is remaining to expand against? That flow is moving fast, and chemistry takes time. How do you make it happen before there is no more nozzle?

Well, you use a fairly long linear aerospike (perhaps with an inflection point for a convex terminus), and you use the combination shockwave to induce combustion as rapidly as possible. The weight cost of the longer aerospike is compensated for by the tremendously high theoretical maximum T/W ratio of your monopropellant thrusters.

This engine would be designed into the airframe of your vehicle so that the extra-long aerospike served as a structural element. The turbopumps and associated machinery/piping would all be inside the aerospike to save space.

[YNM]

Hmm... Yes, I think the problem with your engine is in the "how to impart the momentum ?" and additionally "how to keep the chemical reaction steps correct ?". Other than the resulting combustion might not find any part of your ship to impart momentum, the fire could flow upwards (?), destroying the decomposition on-site... Making it no different than usual HTP-Hydrazine hypergolic engine. Just with aerospike expansion chamber.

Edited February 18, 2016 by YNM

[sevenperforce]

6 minutes ago, YNM said:

Hmm... Yes, I think the problem with your engine is in the "how to impart the momentum ?" and additionally "how to keep the chemical reaction steps correct ?". Other than the resulting combustion might not find any part of your ship to impart momentum, the fire could flow upwards (?), destroying the decomposition on-site... Making it no different than usual HTP-Hydrazine hypergolic engine. Just with aerospike expansion chamber.

I think that the ideal design would have one flow moving down the upper portion of the aerospike in supersonic flow, with the other flow impinging on it at a slight angle. Thus, the standing-wave termination shock would be substantially vertical with respect to the wall of the aerospike. Basically inertial confinement. This allows combustion without choking the flow. The exhaust then travels down the aerospike, expanding against the atmosphere on one side and the aerospike on the other. 

[YNM]

21 minutes ago, sevenperforce said:

I think that the ideal design would have one flow moving down the upper portion of the aerospike in supersonic flow, with the other flow impinging on it at a slight angle. Thus, the standing-wave termination shock would be substantially vertical with respect to the wall of the aerospike. Basically inertial confinement. This allows combustion without choking the flow. The exhaust then travels down the aerospike, expanding against the atmosphere on one side and the aerospike on the other. 

Somebody with access to advanced CFD would need to simulate that...

[shynung]

36 minutes ago, sevenperforce said:

I think that the ideal design would have one flow moving down the upper portion of the aerospike in supersonic flow, with the other flow impinging on it at a slight angle. Thus, the standing-wave termination shock would be substantially vertical with respect to the wall of the aerospike. Basically inertial confinement. This allows combustion without choking the flow. The exhaust then travels down the aerospike, expanding against the atmosphere on one side and the aerospike on the other. 

I'm guessing that the structural part that doubles as the aerospike would be covered in RCC.

[sevenperforce]

3 minutes ago, shynung said:

I'm guessing that the structural part that doubles as the aerospike would be covered in RCC.

Yes, probably. The nozzle would likely be actively cooled, probably by the hydrogen peroxide due to its terrific heat capacity. This is another advantage; the actual combustion chambers get to operate at a much lower temperature because the bulk of combustion takes place against the nozzle, which already requires cooling.

21 minutes ago, YNM said:

Somebody with access to advanced CFD would need to simulate that...

Oh, for sure. But it would be mostly a matter of determining the right parameters to make it work. 

[shynung]

A question, if you don't mind.

Your engine design burns hydrazine as fuel. How much do you think changing the fuel into a hydrazine-derived fuel would do to the performance? Say, switching to MMH, UDMH, or A-50?

Edited February 18, 2016 by shynung

[sevenperforce]

19 minutes ago, shynung said:

A question, if you don't mind.

Your engine design burns hydrazine as fuel. How much do you think changing the fuel into a hydrazine-derived fuel would do to the performance? Say, switching to MMH, UDMH, or A-50?

Good question. I chose hydrazine because it produces nothing but diatomic hydrogen and diatomic nitrogen, but there might be other monopropellants with better performance. The nice thing about hydrazine + peroxide is that both produce completely non-toxic exhaust individually and combust together with a simple, high-impulse H₂ + O₂ reaction. I'd definitely be interested in other possible fuels. Many monopropellants are more like blended/bonded bipropellants, though; for example, nitromethane is unsuited for this application because it contains both the reductant and the oxidizer in itself.

Dinitrogen pentoxide is a relatively unstable salt which can also exist as a covalently-bonded gas; it has terrific oxidizing potential as well as decent monopropellant peformance. I don't know how hard it would be to make it a liquid. Might be easier to simply dissolve dinitrogen pentoxide salt into hydrogen peroxide, if that's possible. I'm better at physics than I am at chemistry. Liquified nitrous oxide is another possibility, though its oxidizing potential by unit mass isn't as great.

The staged-combustion approach is going to give better performance to high-energy-density fuels as opposed to high-specific-energy fuels, because the dynamic pressure is going to depend heavily on the density of the precombustion exhaust streams.