RCAIDE.Library.Methods.Aerodynamics.Common.Lift.BET_calculations.compute_airfoil_aerodynamics

compute_airfoil_aerodynamics(beta, c, r, R, B, Wa, Wt, a, nu, airfoils, airfoil_locations, ctrl_pts, Nr, Na, tc, use_2d_analysis)

Cl, Cdval = compute_airfoil_aerodynamics( beta,c,r,R,B,

Wa,Wt,a,nu, airfoils,a_loc ctrl_pts,Nr,Na,tc,use_2d_analysis )

Computes the aerodynamic forces at sectional blade locations. If airfoil geometry and locations are specified, the forces are computed using the airfoil polar lift and drag surrogates, accounting for the local Reynolds number and local angle of attack.

If the airfoils are not specified, an approximation is used.

Assumptions: N/A

Source: N/A

Inputs:

beta blade twist distribution [-] c chord distribution [-] r radius distribution [-] R tip radius [-] B number of rotor blades [-]

Wa axial velocity [-] Wt tangential velocity [-] a speed of sound [-] nu viscosity [-] airfoil_data Data structure of airfoil polar information [-] ctrl_pts Number of control points [-] Nr Number of radial blade sections [-] Na Number of azimuthal blade stations [-] tc Thickness to chord [-] use_2d_analysis Specifies 2d disc vs. 1d single angle analysis [Boolean]

Outputs:

Cl Lift Coefficients [-] Cdval Drag Coefficients (before scaling) [-] alpha section local angle of attack [rad]