Abstract:

Abstract Combustion and flow thermogasdynamics is one of the most sophisticated computational stages in liquid rocket engine (LRE) design. A considerable amount of different methodologies used to conduct the thermogasdynamic analysis of the combustion and flow process and obtain the accurate estimates of the parameters of the LRE chamber is known, currently, in rocket and space sciences. However, the development of a generalized mathematical model that would have a capability of being applied for any unique combustion case by, hence, yielding reliable results and, nevertheless, be efficient in any application still remains to be an issue. The following paper considers the development of a generalized tensor model the application of which for any desired propellant configuration yields unique results. In addition considered the application of the mentioned model at significant nozzle sections. Moreover the current paper, also, considers an approach based on the specific area in order to compute the parameters at the nozzle throat. The uniqness of this approach is based on the fact, that the outcoming results appear to be more accurate than the ones obtained using convential approaches. The paper accurately summarizes the application of the generalized tensor model based on an example problem which involves the combustion of liquid methane with liquid oxygen.

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