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Integration of airframe and propulsion system is one of the most challenging steps in flight vehicle design cycles. In this paper, a three-dimensional supersonic inlet based on the wave-derived geometry technique has been designed and analyzed. Although the considered method was created for hypersonic forbodies, the idea is fully operational for the low supersonic inlet design at Mach 1.6. The inlet concept in this paper is formed from predefined profile elements which are used to generate the three-dimensional geometry in an oblique shock pattern. By this approach, the curved corner of the inlet entrance edge can generate the same shock as the main compression surface and also these curved surfaces provide the optimum transition between entrance geometry and compressor face which is important for the airflow quality and propulsive efficiency. The new concept has been validated by a series of accurate CFD simulations with completely structural grid domains. The major inlet's performance factors like total pressure recovery, flow distortion and mass flow capture ratio are calculated. The concept and it's accurate numerical simulations create a baseline for more advanced designs and researches about the three-dimensional inlets and geometry transition techniques between the different sections of duct.

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In this paper, a hypersonic inlet for operating at Mach 5.0 is designed and analyzed numerically. The main axis of this study is a series of three-dimensional simulations with the accuracy of 10E-06 which are applied to determine the effects of the highly developed boundary layer on the performance of inlet for three different study cases. The basic inlet concept is designed by integration of double ramp compression surface and inlet duct which can reduce the free-stream Mach number to the range of 2.0. The most important factor that it affects the performance of the hypersonic inlet system, is the developed entropy layer on the fuselage of the flight vehicle. Ingestion of this layer results in thermal gradients and pressure recovery losses. The bow shocks at the nose and the leading edges are the main sources of this low kinetic energy layer. Using the k-ω turbulence model in the numerical simulations have resulted in a reliable estimation of the boundary layer. In the current context, shock structures, shock-boundary layer interactions, flow quality at the end of the diffuser and also the effects of using sidewalls on the performance of the hypersonic inlet are the main goals of the simulations and the related results are summarized