Unmanned aerial vehicles (UAVs) are increasingly utilized in modern warfare, requiring propulsion systems with wide range operating capability and higher speeds to improve multifunctionality and counter threats. Therefore, the demand for a supersonic micro gas turbine engine as a propulsion system for small, high-speed UAVs is increasing, and the requirement for the development of an afterburner for the micro gas turbine engine is growing. So in this study aims to design and evaluate the performance of a diffuser for a micro gas turbine engine-based afterburner. The afterburner was developed based on a commercial turbojet engine. Simulations were conducted using ANSYS CFX, incorporating the k-ω SST turbulence model and the Eddy Dissipation Model for combustion. Fuel behavior was modeled using the Liquid Evaporation Model with Light Oil Modification and Reitz and Diwaker model. Boundary conditions were defined using turbine outlet profiles at a core engine speed of 104,000 RPM. Six diffuser configurations were evaluated by varying tail cone designs and strut geometries (strut angles of 0°, 2.5°, and 5°, with strut thicknesses of 2 mm and 4 mm). The baseline tail cone design has led to recirculation zones within the diffuser, which caused the backflow of combustion gases. To address this issue, the tail cone was redesigned, and the struts were introduced to reduce recirculation zones and decrease swirl components in the flow. The simulation results showed that a strut angle of 2.5° demonstrates the most effective swirl reduction, while maintaining pressure losses below 5% in all the cases. Furthermore, the analysis revealed that strut thickness has a greater impact on diffuser performance than the strut angle.