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The propulsion system integration of a turboprop aircraft, which has been developed for the basic trainer, was performed. The proper turboprop engine was selected among worldwide existing engines by the specific developed engine selection technique and trade‐off studies such as customer’s request for operational capability (ROC), propulsion system parameters, performance analysis with engine installed effects, future growth potential, integrated logistic support (ILS), maintainability, interfaces with the airframe, etc. The chin type air inlet with the plenum chamber was designed in consideration of the inclined configuration to minimize the propeller swirl effect, the inertial separation bypass device to reduce FOD, and the super‐ellipse and NACA‐1 profile lip to maximize the ram recovery. The air inlet was analyzed by a higher‐order source panel method considering propeller wake. The exhaust duct was designed through internal cross‐section area determination to maximize the cruising power as well as external configuration to maximize the effective power, to minimize the aerodynamic drag and to minimize the cockpit contamination by the exhaust gas. The proper oil cooler for the selected turboprop engine was determined with cooling requirements and the oil cooling inlet duct with NACA configuration was designed. The test of the propulsion system including the installation performance test with the effects of the air inlet, the exhaust duct, the propeller and the nose fuselage configuration was performed in the test cell.

To set‐up a specific design procedure for the smart unmanned aerial vehicle (UAV) fuel supply system which has been developed by Korean Aerospace Research Institute, and to design it preliminarily with the fuel system requirement and target reliabilities.

The fuel system layout and fuel tank were determined through consideration of total fuel volume, fuel flow rate, reliability, weight, centre of gravity, etc. In sizing of components such as booster pumps, jet pumps, piping system, vent subsystem, refuelling and defuelling subsystem, engine fuel flow requirement, pressure loss, component failure rate, weight and centre of gravity were considered. Finally, the reliability analysis of the preliminary designed fuel system was carried out.

According to the reliability analysis and weight estimation results, it was confirmed that the proposed fuel system agreed well with the design specifications and target reliabilities required by the vehicle system.

In current preliminary design phase, the most important consideration is the reliability of the fuel system. Therefore, the weight estimation of the designed fuel system to meet this reliability requirement could not meet partially the system's requirements. In the next design step, the proper fuel system for weight reduction will be performed through an optimization process between weight and reliability.

A specific design procedure components' sizing to meet system requirement target reliability for UAV vertical take‐off/landing was proposed.