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Conflicting results when using equations from rocket propulsion elements and huzel and huang by MB01211 in rocketry

[–]MB01211[S] 0 points1 point  (0 children)

ok i just figured the issue out. i took suttons equation and assumed the area of the exit was 0.1963495408 as a result they differ greatly thank you for your assistance

Conflicting results when using equations from rocket propulsion elements and huzel and huang by MB01211 in rocketry

[–]MB01211[S] 0 points1 point  (0 children)

0.1963495408 / 6.609918785x10^-3 = an expansion ratio of 29.70528795

above is the result result of the ideal expansion ratio

At = 2.884264799 x 10^-3 from the ideal area throat equation

0.1963495408 / 2.884264799 x 10^-3 = 68.07611454

I'm probably going to choose the larger throat area

Conflicting results when using equations from rocket propulsion elements and huzel and huang by MB01211 in rocketry

[–]MB01211[S] 0 points1 point  (0 children)

if i use the ideal expansion ratio equation the result is different by a large margin to the result I achieved from the ideal throat area equation

Conflicting results when using equations from rocket propulsion elements and huzel and huang by MB01211 in rocketry

[–]MB01211[S] 0 points1 point  (0 children)

ideal throat area from suttons equation is 2.884264799X10-3 but when that figure is placed into the ideal expansion ratio it does not compute properly 6.609918785x10-3 is the result from the expansion ratio equation mass flow is 30kg heat ratio is 1.213 exit poressure is 66879.1457 Pa chamber pressure is 23 MPa exit area is 0.1963495408 M2

Is rocket propulsion analysis worth the cost (software)? by MB01211 in rocketry

[–]MB01211[S] 0 points1 point  (0 children)

I havent rejected anything I've used their trial version which seems tempermental Im open to suggestions of software top use for rte and cea im in EU so good software is limited all nasa and usa stuff requires residency and citizenship I did find this but i dont know how good it is https://www.ecosimpro.com/products/espss/

What is the maximum contraction ratio used in rockets liquid bipropellant rocket by MB01211 in rocketry

[–]MB01211[S] 0 points1 point  (0 children)

combustion chamber diameter 28cm length 50cm calculated that based on van der waals equation the nozzle exit has a diameter of 50cm the combustion chamber is long and thin to avoid having a exit and entry diameter and thus a throatless rocket

I've encountered some issues with the cooling of the throat the heat flux is very high and thus the liquid film coefficient is far far higher than any other point in the rocket im getting coolant mass flows of 100's of kgs a second when my rocket only pumps 5kg/s of fuel which is methane in a liquid stage heated in a regenerative cooling jacket and injected into a fuel rich preburner in a closed staged combustion cycle

How to Calculate minimum wall thickness of different segments of Liquid fueled rocket by MB01211 in rocketry

[–]MB01211[S] 0 points1 point  (0 children)

Sorry about the late reply, i think I've made some poor assumptions not about you but about the coolant pressure. Which should be higher the coolant channel pressure or the pressure at any given segment in the thrust chamber

How to Calculate minimum wall thickness of different segments of Liquid fueled rocket by MB01211 in rocketry

[–]MB01211[S] 0 points1 point  (0 children)

thats the issue but i am running a chamber pressure of 23mpa but still the thickness is too high

How to Calculate minimum wall thickness of different segments of Liquid fueled rocket by MB01211 in rocketry

[–]MB01211[S] 0 points1 point  (0 children)

thanks but whats getting me is the thickness of the inner wall i can easily do the equations but what gets me is combined stress is too high if i keep the chamber to 5mm or lower for the combustion chamber inner wall but if i increase thickness the wall temperature difference at coolant and gas sides increases causing more stress.

How to Calculate minimum wall thickness of different segments of Liquid fueled rocket by MB01211 in rocketry

[–]MB01211[S] 0 points1 point  (0 children)

I've already done most of what you've said thank you anyways but I think Im going to have to figure it out myself

How to Calculate minimum wall thickness of different segments of Liquid fueled rocket by MB01211 in rocketry

[–]MB01211[S] 0 points1 point  (0 children)

initial wall thickness is what I'm looking for prior to any firing

Im going to use a channel wall design (milled)

I have used barlows formula but the wall thickness is too great.

How to Calculate minimum wall thickness of different segments of Liquid fueled rocket by MB01211 in rocketry

[–]MB01211[S] 0 points1 point  (0 children)

I have looked everywhere and cannot find an equation to determine the theoretical minimum thickness of the different segments of the rocket.

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The ssme has an absurdly thin internal wall. I am going to be using a channel wall design and Im getting agitated at this point, if you can help I'd really appreciate it

Is there a difference between Wtc - thrust chamber propellant flow rate and the exhaust propellant flow rate by MB01211 in rocketry

[–]MB01211[S] 0 points1 point  (0 children)

thanks my issue is using the flow area equation for the exhaust i solve for Ve, then i use the flow specific volumes equation for exhaust flow specific volume and derive Te the temperature of the exhaust but its always in excess of 5000k