In this study, an air intake was designed for the combustion chamber of a supersonic projectile operating at flight conditions of Mach 3.4 and an altitude of 15 km. Initially, a three-dimensional design methodology for the supersonic air intake is presented. Subsequently, the designed intake was integrated with the combustion chamber, and its performance was evaluated. To ensure the accuracy of the analyses, the simulation process was validated against the results of an existing combustion chamber. The integrated geometry was then meshed and analyzed. The results indicate that the air intake’s performance closely aligns with the calculated theoretical values. The maximum error observed was 6.25%, corresponding to the Mach number at the first section of the intake. The obtained total pressure recovery factor also showed a 2.43% error compared to the calculated values. The performance of the combustion chamber, using the airflow supplied by the intake, was examined, yielding a combustion temperature of 1298 K and a combustion efficiency of 83.5%. Furthermore, the distance between the fuel injection point and the air inlet was reduced to investigate its effect on combustion. It was determined that at a distance-to-diameter ratio of 1, the combustion efficiency increased by 3.9%.