A Fuzzy Velociraptor

Empty weight of solid stages tend to be lower than empty weight of liquid engines. Solid enignes tend to have a lower ISP than liquid engines which is why they require more stages to get to orbit.

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.

The engine would break and the rocket would either fall down hit the ground and explode or have to be blown up by launch abort. The F1 engine is a LOX/RP-1 engine. The SII (second stage of saturn V) used LOX/LH2. LH2 behaves very differently from RP1 in temperature, density, mass flow rate etc. While there are some engines that a fuel or oxidizer can be substituted RP-1 switching with LH2 is almost certainly not one of those cases.

Titan's pretty neat but I think answering the poll literally Triton is the coolest moon.

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.

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.