###### Abstract  
Due to the extreme temperatures inside the combustion chambers of liquid propellant rocket engines, the walls of the combustion chamber and the nozzle are cooled by either the fuel or the oxidizer in what is known as _regenerative cooling_. This study presents an indigenous computational tool developed for the analysis of heat transfer in regenerative cooling of such rocket engines. The developed tool incorporates a one-dimensional (1-D) combustion analysis to calculate the thermophysical properties of the combustion gas. Basic engine properties were calculated and used to generate a thrust chamber profile based on a bell-shaped nozzle. The hot gas side was analyzed using 1-D isentropic flow assumptions, along with heat transfer correlations. The coolant side was evaluated using the hydraulic analysis in the axial direction and the heat transfer analysis in the radial direction. Thermophysical properties and the phase of the coolant were determined using the given property tables and the instantaneous state of the coolant. This flexible and computationally less demanding tool was used to analyze two small-scale engines utilizing liquid hydrocarbon fuels, which are used in modern rocket propulsion. The wall cooling analyses of a liquid oxygen (LOX)/liquid methane (LCH\({}_{4}\)) engine and a liquid oxygen (LOX)/liquid propane (LC\({}_{3}\)H\({}_{8}\)) engine are presented. Fuel and oxidizer were used separately as coolants for both engines, and both of them experienced phase change. Results reveal the advantage of the high mass flow rate of the oxidizer in cooling performance. In addition, the results of this study show that the cooling of the LOX/LC\({}_{3}\)H\({}_{8}\) engine is somewhat more challenging compared to the LOX/LCH\({}_{4}\) engine. In the early design phase, engineers may desire to assess whether the selected fuel is sufficient for regenerative cooling in terms of thermophysical properties and mass flow rate, or whether an additional cooling method such as film cooling should be combined with regenerative cooling. In addition, occasionally, oxidizer could be preferred as the coolant, as in the case of some liquid oxygen (LOX)/kerosene engines [3]. Furthermore, in the early design phase, one could desire to have preliminary information about the size and the shape of the cooling channels, as it influences the overall design and manufacturing. As the development process progresses, detailed analyses regarding the engine cooling will have to be conducted; however, the early decisions should be on time and logical, so as to not result in any inconvenience or lost time in ensuing phases. Therefore, simple modelling tools for regenerative cooling of liquid propellant rocket engines are invaluable in early design phases. Such tools could be used to evaluate the suitability of regenerative cooling, and to find the optimum channel number and cross section as a starting point. Heat transfer in liquid propellant rocket engines has been extensively studied in the literature. RTE [4] and TDK [5] are two well-known tools for the thermal analysis of rocket engines that are developed using older programming methods. Despite the utilization of these tools in different studies [6; 7; 8], simpler and more practical modelling approaches are preferred in some simple liquid propellant rocket engine thermal analysis studies. For instance, [PERSON] et al. [9] presented a 1-D modelling approach for regeneratively cooled rocket engines. In that study, a thermal resistance network approach was used by treating the cooling channel walls as fins. The authors compared their results with the engine temperature measurements of the Space Shuttle Main Engine (SSME) [10]. [PERSON] et al. [11] established a program which, based on given engine design requirements, creates a thrust chamber profile and carries out a preliminary 1-D regenerative cooling analysis. They focused on LOX/kerosene engines with this methodology. [PERSON] et al.