--------------------------------------------- Run case 1: level flight, Xcg=3.0 alpha -> CL = 0.600000 beta -> beta = 0.00000 pb/2V -> pb/2V = 0.00000 qc/2V -> qc/2V = 0.00000 rb/2V -> rb/2V = 0.00000 camber -> camber = 0.00000 aileron -> Cl roll mom = 0.00000 elevator -> Cm pitchmom = 0.00000 rudder -> Cn yaw mom = 0.00000 alpha = 2.13442 deg beta = 0.00000 deg pb/2V = 0.338044E-18 qc/2V = 0.00000 rb/2V = 0.730949E-19 CL = 0.600000 CDo = 0.200000E-01 bank = 0.00000 deg elevation = 0.00000 deg heading = 0.00000 deg Mach = 0.00000 velocity = 5.42671 m/s density = 1.22500 kg/m^3 grav.acc. = 9.81000 m/s^2 turn_rad. = 0.00000 m load_fac. = 1.00000 X_cg = 3.00000 m Y_cg = 0.00000 m Z_cg = 0.609524 m mass = 0.231000 kg Ixx = 0.165803E-01 kg-m^2 Iyy = 0.113692E-01 kg-m^2 Izz = 0.278108E-01 kg-m^2 Ixy = 0.304560E-10 kg-m^2 Iyz = -0.135360E-10 kg-m^2 Izx = -0.362168E-03 kg-m^2 visc CL_a = 0.00000 visc CL_u = 0.00000 visc CM_a = 0.00000 visc CM_u = 0.00000