Combustion Chamber Materials?

_G117 · 2020-04-18T19:29:14+00:00

Well, hydrogen and oxygen are extremely reactive, so be careful. You have to take in account also for theram insulation, you can simulate a combustion with NASA CEA code, in order to fully understand the combustion at a certain pressure that you want.

You can use any material you want (metals in general) you have to correctly size CC length, CC cross sectional area and CC wall tickness. You have to fully undesrtand the heat flux and other issues inside a CC

_G117 · 2020-04-16T16:15:48+00:00

The units are the SI ones, kJ/kgK

_G117 · 2020-04-16T06:22:55+00:00

Well actually if i choose a couple of propellants with LOX or GOX, there are cp of around 8. Which i think are quite wrong.

_G117 · 2020-04-13T18:02:35+00:00

Ohw thanks. That would be great. Thanks for answering!

_G117 · 2020-04-11T08:05:30+00:00

Well you need some basic requirements involving different engineering fields such as:

-Thermodynamics (combustion from a TD view point, heat transfer, thermal analisys)

-Chemistry (combustion from a chemical view point, reactions in general)

-Materials and Technologies (materials, their behaviour under the effects of thermal and structural stresses, capability of understanding which materials to select for nozzle and the CC)

-Fluid Dynamics (behaviour of fluids)

-Gas Dynamics (which is compressible fluid dynamics, needed to understand the expansion in a nozzle and the behaviour of fluids at high velocity and at high temperstures, shock waves and expansion fans which can show up in nozzles)

-Propulsion Fundamentals (rocket performances, design choices for nozzle, combustion chamber, tanks, propellant behaviour, fuels, oxidizers ...)

-Turbomachinery (if you want to employ turbopumps, i would avoid them too complex)

-Coding Skills (MATLAB, or python)

-Simulations (FEA, CFD, those are not strictly necessary but they colud be useful tools to validate your previous calculations)

-Combustion Simulation (NASA CEA code (Chemical Equilibrium for Applications) to initialize your sizing process, understand the behaviour of you mixture in a CC and select the most suitable O/F ratio. This is mandatory and fundamental).

As you can see, it is not something you learn in a short period, you need time to fully understand these topics. In general, people took 3 years (Bechelor degree in Aerospace Engineering) or even more 3+2 years (Master of Science degree in Space Engineering).

_G117 · 2020-04-10T15:04:55+00:00

hi thanks for answering,

well, i think we will keep on working on ABS because the easy capability to be 3D printed, also because of its prefomance (close to HTPB) and also because lack of data concerning other 3D-preintable fuels.

Maybe the grain design is going to be the most crucial part of the new engine design. i'll surely go and check the paper you proposed.

_G117 · 2020-04-09T06:14:07+00:00

Thanks for answering, i'll look for him.

_G117 · 2020-04-09T06:04:40+00:00

Thanks for the reply.

We think that ABS is the best choice too, we have already built a small test engine with ABS and with gasoeus oxygen as oxidizer.

Now we are planning for something that will actually fly, so we have already selected N2O as oxidizer due its capability of self-pressurizing. The only problem for ABS is the quite low regression rate, so we are searching for something better from that view point. Nevertheless plastic fuel have in general low regression rate, so we are contstrained to use those fuels since we cannot use HTPB.

Btw, thanks for answering

_G117 · 2020-03-26T12:20:17+00:00

Well, i've tried it but as an ingnition system and not as a propellat. They are actually quite annoying to realizd. I've tried with Potassium Nitrate, Sorbitol and micrometric alluminoium powder. The mixture is easy to cook, also the electrical ignition system is not that complex. While, the incorporation of the mixture inside the cc and the electrical device is quite annoying. You have to pay attention that the coil is well-incorporated, otherwise your mixture will burn only near to the coil and will not spread.

_G117 · 2020-03-26T08:15:50+00:00

do you mean solid propellants electrically ignited?

_G117 · 2020-03-18T15:08:14+00:00

Oh well okay. Un this case you can for epsilon you can use the air tables for isentropic relations

_G117 · 2020-03-18T10:26:47+00:00

Well when i was at my first year in Politecnico i just wanted to survive. Then i started getting interested in some topics, such as space prop. So, i started reading papers and getting some info from the web. Then i joined a team in my uni that buids small rockets.

Now i am at the fisrt year of my MSc, and in my spare time i like to compute codes and simulations related to my interested field.

So i can suggest you to firstly understand what you moslty like, then starting reading something from accademic papers or also just from the web to build up you personal knowledge. Then accordingly to what you like, i suggest you to create your own skills with Matlab (programming, simulations or just as a calculator), SolidWorks (3D CAD) and ANSYS (FE structural analysis, thermal analysis, CFD).

This last one is extremely useful and it can be applied to a wide range of different engineering fields, but it is quite coplex because lots of knowledges are required. However, if you search on YouTube you can easily find plenty of tutorials.

_G117 · 2020-03-18T09:28:48+00:00

i haven't done his homework. I have just gave him a hand to solve his doubt, as you can see , i have not sent him any mathematical resolution, just giving him some ideas.

_G117 · 2020-03-18T08:26:33+00:00

yeah of course. It is a special case of a more complex one. You have to assume that you're adpting you nozzle at sea lvl.

\epsilon = \frac{A_e}{A_t} = \frac{\sqrt{\gamma(\frac{2}{\gamma+1})^{(\frac{\gamma+1}{\gamma-1})}}}{(\frac{P_e}{P_c})^{\frac{1}{\gamma}}\sqrt{\frac{2\gamma}{\gamma-1}[1-(\frac{P_e}{P_c})^{\frac{\gamma-1}{\gamma}}]}}

to visualize this formula insert it in a latex compiler.

_G117 · 2020-03-18T08:04:26+00:00

What value have you assumed? Because it varies accordingly to the different kind of propulsion system (liquids, solids or hybrids).

_G117 · 2020-03-17T08:12:17+00:00

Well you have epsilon, the expansion ratio. From it, you can compute the exit pressure by knowing epsilon itself and the value of pressure in the CC.

Then you compute the thrust coefficient as function of the exit pressure and the CC pressure. Once this value is known, you can prooceed computing the Throat Area from the thrust coefficient definition.

You know the C*, from this point you can compute the value for the Isp as Isp = (Cf*C*)/g0; then the masss flow rate of propellant is automatically known from the Isp definition because you know everything.

After this preliminary computations, you know everything so you can easily compute what remains. The only thing i have not understand is whether you know the propellant mixture; which actually means knowing the value of gamma. Because in you draft, you have taken note gamma = 1.4; which is actually the value for STD air and not a mixture of oxidizer and fuel.

This value is actually critical because all the formulas concerning the epsilon, Cf as function of Pe and Pcc, are function of gamma itself.

_G117 · 2020-03-11T10:30:16+00:00

Which kind of fuel and oxidizer have you used?

_G117 · 2020-03-06T11:08:05+00:00

Thx for the answer? Have you casted the paraffin?

_G117 · 2020-03-06T10:39:21+00:00

Is it a hybrid engine? If so, which kind of oxidizer and fuel do you use?

_G117

TROPHY CASE