Abstract
Accurate determination of heat flux is an important task not only in the designing aspect, but also in the performance analysis of rocket engines. In this purpose, this work deals with the heat flux determination in a combustion chamber through the inverse method. In this approach, the transient heat flux is determined from the experimental temperature data measured at the outer sidewall of the rocket engine. In this work the physical phenomenon was modeled by the transient one-dimensional heat equation in cylindrical coordinates and the material properties of the chamber were considered constant. Furthermore, the model is solved using the inverse heat conduction problem with least squares modified by the addition of Tikhonov regularization term of zero-order. Moreover, the sensitivity coefficients were obtained by Duhamel’s theorem. Through the regularization parameter, it was able to generate acceptable results even when using data with considerable experimental errors.
Highlights
In a thrust chamber, the amount of energy transferred as heat to the chamber walls is between 0.5% and 5% of the total generated energy (Sutton 1992)
Theoretical and Experimental Heat Transfer in Solid Propellant Rocket Engine 1x7x/x1x8 where Ts is the temperature of combustion chamber before test (19 °C); and is the temperature of the hot gases (1302.47 °C), given by Propep software
A program code was developed for solving Inverse Problem (IP) generated results, which were compared to the results accomplished through analytical correlations from the literature
Summary
In a thrust chamber (nozzle and combustion chamber), the amount of energy transferred as heat to the chamber walls is between 0.5% and 5% of the total generated energy (Sutton 1992). This amount could be enough to cause structural failure. The computation of the radiation heat transfer must be accounted by the surfaces emissivity and the absorption and scattering coefficients of the fluid mixture. Information about these parameters is not always found. Considering the propellant applied in this work (potassium nitrate with sucrose, KNSu), the combustion products that affects radiation heat transfer are predominantly
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