Aerodynamic Design of the NASA Rotor 67 for Non Uniform Inflow

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Aerodynamic Design of the NASA Rotor 67 for Non Uniform Inflow ( aerodynamic-design-nasa-rotor-67-non-uniform-inflow )

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Master Thesis Report 3D Flow Simulation of the NASA Rotor 67 Spanwise Rotor Total Pressure Upstream Total Press Downstream Total Pr ure essure 25 20 15 10 5 90000 100000 110000 120000 130000 140000 150000 160000 170000 Pressure [Pascal] Spanwise Rotor Total Temperature Upstream Total Temperature Downstream Total Temperature 25 20 15 10 5 280 290 300 310 Temperature [Kelvin] 320 330 340 350 Figure 7.1: Spanwise Total Pressure at Inlet/Outlet Boundary Figure 7.2: Spanwise Total Temperature at Inlet/Outlet Boundary Spanwise Rotor Static Pressure Upstream Static Pressure Downstream Static Pressure 25 20 15 10 5 80000 90000 100000 110000 120000 Pressure [Pascal] 130000 Figure 7.3: Spanwise Static Pressure at Inlet/Outlet Boundary Figure 7.4: Measurement Points at Inlet and Outlet Boundary An illustration of the measurement points taken at both the inlet and outlet boundary condition is shown in figure 7.4. Using this boundary condition, the total to static pressure ratio at the mid radius was obtained. A value of approximately 1.2 was obtained. 7.2 3D Non Uniform Inflow Boundary Conditions In order to produce a non uniform inflow boundary condition, a previous study on the design of op- timum inlet configuration for BLI was used as a reference. Since the stagnation pressure loss across the inlet is non-uniform due to the fact that fuselage boundary layer exists primarily on the bottom segment of the aerodynamic interface plane, the circumferential averaged total pressure at the aero- 72 Spanwise Radial Position [cm] Spanwise Radial Position [cm] Spanwise Radial Position [cm]

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