A theoretical study analyzing three-dimensional combustion acoustic instabilities in a liquid PROPELLANT rocket engine combustor has been conducted. A linear theory based on Crocco's pressure sensitive time lag model is used. To apply this theory, the combustor is divided into two main components, including the combustion chamber and the converging part of the nozzle. The assumption of concentrated combustion zone is used and the governing perturbation equations describing oscillations of flow variables are considered. To solve these equations appropriate boundary conditions at both ends of the combustion chamber are req1lired. Combustion zone boundary condition at one end and the nozzle admittance relation at other end are used. To obtain the nozzle admittance the three dimensional flow perturbation equations are solved in the converging part of the nozzle. This approach is capable of predicting acoustic stability behavior of a combustor at a wide range of Mach numbers and frequencies. Also, this analysis enables the rocket engine designer to observe the effects of different parameters such as nozzle entrance Mach number, chamber geometry, nozzle geometry. and gas properties on stability characteristics of an engine combustor. In case of instability observation; one can predict the acoustic mode which causes the instability and achieve an optimum design before conducting any expensive and time consuming experimental tests. This paper presents the stability analysis results and a parametric study of the effect of design parameters on stability characteristics of a typical combustor.