--------------------------------------------- Run case 1: 0 deg bank alpha -> CL = 0.700000 beta -> beta = 0.00000 pb/2V -> pb/2V = 0.00000 qc/2V -> qc/2V = 0.00000 rb/2V -> rb/2V = 0.00000 flap -> flap = 0.00000 aileron -> Cl roll mom = 0.00000 elevator -> Cm pitchmom = 0.00000 rudder -> Cn yaw mom = 0.00000 alpha = 3.69469 deg beta = 0.00000 deg pb/2V = 0.180670E-43 qc/2V = 0.00000 rb/2V = 0.905603E-44 CL = 0.700000 CDo = 0.150000E-01 bank = 0.00000 deg elevation = 0.00000 deg heading = 0.00000 deg Mach = 0.00000 velocity = 6.54829 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 = 3.75000 Y_cg = 0.00000 Z_cg = 0.00000 mass = 1.25020 kg Ixx = 0.474178 kg-m^2 Iyy = 0.957865E-01 kg-m^2 Izz = 0.566328 kg-m^2 Ixy = 0.690336E-09 kg-m^2 Iyz = 0.00000 kg-m^2 Izx = -0.265910E-02 kg-m^2 visc CL_a = 0.00000 visc CL_u = 0.00000 visc CM_a = 0.00000 visc CM_u = 0.00000 --------------------------------------------- Run case 2: 0 deg bank alpha -> CL = 0.500000 beta -> beta = 0.00000 pb/2V -> pb/2V = 0.00000 qc/2V -> qc/2V = 0.00000 rb/2V -> rb/2V = 0.00000 flap -> flap = 0.00000 aileron -> Cl roll mom = 0.00000 elevator -> Cm pitchmom = 0.00000 rudder -> Cn yaw mom = 0.00000 alpha = 1.69687 deg beta = 0.00000 deg pb/2V = -0.115570E-32 qc/2V = 0.00000 rb/2V = -0.797920E-33 CL = 0.500000 CDo = 0.150000E-01 bank = 0.00000 deg elevation = 0.00000 deg heading = 0.00000 deg Mach = 0.00000 velocity = 7.74805 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 = 3.75000 Y_cg = 0.00000 Z_cg = 0.00000 mass = 1.25020 kg Ixx = 0.474178 kg-m^2 Iyy = 0.957865E-01 kg-m^2 Izz = 0.566328 kg-m^2 Ixy = 0.690336E-09 kg-m^2 Iyz = 0.00000 kg-m^2 Izx = -0.265910E-02 kg-m^2 visc CL_a = 0.00000 visc CL_u = 0.00000 visc CM_a = 0.00000 visc CM_u = 0.00000 --------------------------------------------- Run case 3: 0 deg bank alpha -> CL = 0.300000 beta -> beta = 0.00000 pb/2V -> pb/2V = 0.00000 qc/2V -> qc/2V = 0.00000 rb/2V -> rb/2V = 0.00000 flap -> flap = 0.00000 aileron -> Cl roll mom = 0.00000 elevator -> Cm pitchmom = 0.00000 rudder -> Cn yaw mom = 0.00000 alpha = -0.287216 deg beta = 0.00000 deg pb/2V = -0.315713E-33 qc/2V = 0.00000 rb/2V = -0.797293E-33 CL = 0.300000 CDo = 0.150000E-01 bank = 0.00000 deg elevation = 0.00000 deg heading = 0.00000 deg Mach = 0.00000 velocity = 10.0027 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 = 3.75000 Y_cg = 0.00000 Z_cg = 0.00000 mass = 1.25020 kg Ixx = 0.474178 kg-m^2 Iyy = 0.957865E-01 kg-m^2 Izz = 0.566328 kg-m^2 Ixy = 0.690336E-09 kg-m^2 Iyz = 0.00000 kg-m^2 Izx = -0.265910E-02 kg-m^2 visc CL_a = 0.00000 visc CL_u = 0.00000 visc CM_a = 0.00000 visc CM_u = 0.00000 --------------------------------------------- Run case 4: 0 deg bank alpha -> CL = 0.200000 beta -> beta = 0.00000 pb/2V -> pb/2V = 0.00000 qc/2V -> qc/2V = 0.00000 rb/2V -> rb/2V = 0.00000 flap -> flap = 0.00000 aileron -> Cl roll mom = 0.00000 elevator -> Cm pitchmom = 0.00000 rudder -> Cn yaw mom = 0.00000 alpha = -1.27540 deg beta = 0.00000 deg pb/2V = 0.208867E-33 qc/2V = 0.00000 rb/2V = -0.230950E-32 CL = 0.200000 CDo = 0.150000E-01 bank = 0.00000 deg elevation = 0.00000 deg heading = 0.00000 deg Mach = 0.00000 velocity = 12.2507 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 = 3.75000 Y_cg = 0.00000 Z_cg = 0.00000 mass = 1.25020 kg Ixx = 0.474178 kg-m^2 Iyy = 0.957865E-01 kg-m^2 Izz = 0.566328 kg-m^2 Ixy = 0.690336E-09 kg-m^2 Iyz = 0.00000 kg-m^2 Izx = -0.265910E-02 kg-m^2 visc CL_a = 0.00000 visc CL_u = 0.00000 visc CM_a = 0.00000 visc CM_u = 0.00000 --------------------------------------------- Run case 5: 0 deg bank alpha -> CL = 0.150000 beta -> beta = 0.00000 pb/2V -> pb/2V = 0.00000 qc/2V -> qc/2V = 0.00000 rb/2V -> rb/2V = 0.00000 flap -> flap = 0.00000 aileron -> Cl roll mom = 0.00000 elevator -> Cm pitchmom = 0.00000 rudder -> Cn yaw mom = 0.00000 alpha = -1.76875 deg beta = 0.00000 deg pb/2V = 0.312486E-34 qc/2V = 0.00000 rb/2V = 0.126195E-32 CL = 0.150000 CDo = 0.150000E-01 bank = 0.00000 deg elevation = 0.00000 deg heading = 0.00000 deg Mach = 0.00000 velocity = 14.1459 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 = 3.75000 Y_cg = 0.00000 Z_cg = 0.00000 mass = 1.25020 kg Ixx = 0.474178 kg-m^2 Iyy = 0.957865E-01 kg-m^2 Izz = 0.566328 kg-m^2 Ixy = 0.690336E-09 kg-m^2 Iyz = 0.00000 kg-m^2 Izx = -0.265910E-02 kg-m^2 visc CL_a = 0.00000 visc CL_u = 0.00000 visc CM_a = 0.00000 visc CM_u = 0.00000 --------------------------------------------- Run case 6: 0 deg bank alpha -> CL = 0.100000 beta -> beta = 0.00000 pb/2V -> pb/2V = 0.00000 qc/2V -> qc/2V = 0.00000 rb/2V -> rb/2V = 0.00000 flap -> flap = 0.00000 aileron -> Cl roll mom = 0.00000 elevator -> Cm pitchmom = 0.00000 rudder -> Cn yaw mom = 0.00000 alpha = -2.26167 deg beta = 0.00000 deg pb/2V = 0.257107E-33 qc/2V = 0.00000 rb/2V = -0.785999E-33 CL = 0.100000 CDo = 0.150000E-01 bank = 0.00000 deg elevation = 0.00000 deg heading = 0.00000 deg Mach = 0.00000 velocity = 17.3252 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 = 3.75000 Y_cg = 0.00000 Z_cg = 0.00000 mass = 1.25020 kg Ixx = 0.474178 kg-m^2 Iyy = 0.957865E-01 kg-m^2 Izz = 0.566328 kg-m^2 Ixy = 0.690336E-09 kg-m^2 Iyz = 0.00000 kg-m^2 Izx = -0.265910E-02 kg-m^2 visc CL_a = 0.00000 visc CL_u = 0.00000 visc CM_a = 0.00000 visc CM_u = 0.00000