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Airfoil CL & CD Start at normal magnitude then skyrocket. |
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October 26, 2020, 20:48 |
Airfoil CL & CD Start at normal magnitude then skyrocket.
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Trevor T
Join Date: Oct 2020
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I'm working with a few airfoils, NACA 23012, NACA 63015A, NACA 4412, RC4-10, and Boeing VR-7, analyzing the lift, drag and moment of each.
Currently working on the NACA 63015A. Using velocities(mph) with their respective lift/drag/moment coefficients are below.
After 25mph both lift and drag are too high. I recently changed to reference values to get the 25mph to be in its normal range. Airfoils' length are 8.9in = .22606m Here are the details for the 2D case: Solver: Pressure-Based Steady state Velocity formulation -> Absolute Model: Viscous Model -> Spalart-Allmaras (1eqn) Vorticity-Based Fluid: Air Density = 1.225 (kg/m^3) Viscosity = 1.802e-5 (Kg/m-s) Reference Values changes: Area = 0.22606 Density = 1.225 Depth = 1 enthalpy = 0 Length = 0.22606 Pressure = 0 Temperature 288.16 Velocity = varies Viscosity = 1.802e-5 Ratio of specific heat = 1.4 Yplus for heat tran. coef. = 300 Turbulence Specification Method -> Turbulent Viscosity Ratio Ratio = 1 Outlet pressure = 0 Solution Method Coupled Gradient -> Least Squares Cell Based Pressure->Second Order Momentum -> Second Order Upwind Modified Turbulent Viscosity -> Second Order Upwind Pseudo Transient ->off Report definitions of drag x=cos(angle) y=sin(angle) Report definitions of lift x=-sin(angle) y=cos(angle) Report definitions of moment Center x=1/4 of span y= location of x in y direction. ONE LAST THING The contour graphics are showing the boundaries of the elements instead of a filled in contoured graph. Last edited by T-Top; October 27, 2020 at 08:29. |
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airfoil 2d, drag and lift, graphics window |
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