69638edca8afc4fef18c77c78171a808a7d359d6,trunk/SUAVE/Methods/Performance/estimate_take_off_field_length.py,,estimate_take_off_field_length,#Any#Any#Any#Any#,31

Before Change

    // ==============================================
    // Getting engine thrust
    // ==============================================    
    state = Data()
    state.conditions = Aerodynamics() 
    state.numerics   = Numerics()
    conditions = state.conditions

    conditions.freestream.dynamic_pressure = np.array(np.atleast_1d(0.5 * rho * speed_for_thrust**2))
    conditions.freestream.gravity          = np.array([np.atleast_1d(sea_level_gravity)])
    conditions.freestream.velocity         = np.array(np.atleast_1d(speed_for_thrust))
    conditions.freestream.mach_number      = np.array(np.atleast_1d(speed_for_thrust/ a))
    conditions.freestream.speed_of_sound   = np.array(a)
    conditions.freestream.temperature      = np.array(np.atleast_1d(T))
    conditions.freestream.pressure         = np.array(np.atleast_1d(p))
    conditions.propulsion.throttle         = np.array(np.atleast_1d(1.))
    
    results = vehicle.propulsors.evaluate_thrust(state) // total thrust
    
    thrust = results.thrust_force_vector

    // ==============================================
    // Calculate takeoff distance
    // ==============================================

    // Defining takeoff distance equations coefficients
    try:
        takeoff_constants = vehicle.takeoff_constants // user defined
    except:  // default values
        takeoff_constants = np.zeros(3)
        if engine_number == 2:
            takeoff_constants[0] = 857.4
            takeoff_constants[1] =   2.476
            takeoff_constants[2] =   0.00014
        elif engine_number == 3:
            takeoff_constants[0] = 667.9
            takeoff_constants[1] =   2.343
            takeoff_constants[2] =   0.000093
        elif engine_number == 4:
            takeoff_constants[0] = 486.7
            takeoff_constants[1] =   2.282
            takeoff_constants[2] =   0.0000705
        elif engine_number >  4:
            takeoff_constants[0] = 486.7
            takeoff_constants[1] =   2.282
            takeoff_constants[2] =   0.0000705
            print("The vehicle has more than 4 engines. Using 4 engine correlation. Result may not be correct.")
        else:
            takeoff_constants[0] = 857.4
            takeoff_constants[1] =   2.476
            takeoff_constants[2] =   0.00014
            print("Incorrect number of engines: {0:.1f}. Using twin engine correlation.".format(engine_number))

    // Define takeoff index   (V2^2 / (T/W)
    takeoff_index = V2_speed**2. / (thrust[0][0] / weight)
    // Calculating takeoff field length
    takeoff_field_length = 0.
    for idx,constant in enumerate(takeoff_constants):
        takeoff_field_length += constant * takeoff_index**idx
    takeoff_field_length = takeoff_field_length * Units.ft
    
    // calculating second segment climb gradient, if required by user input
    if compute_2nd_seg_climb:
        // Getting engine thrust at V2 (update only speed related conditions)
        state.conditions.freestream.dynamic_pressure = np.array(np.atleast_1d(0.5 * rho * V2_speed**2))
        state.conditions.freestream.velocity         = np.array(np.atleast_1d(V2_speed))
        state.conditions.freestream.mach_number      = np.array(np.atleast_1d(V2_speed/ a))
        results = vehicle.propulsors["turbofan"].engine_out(state)
        thrust = results.thrust_force_vector[0][0]

        // Compute windmilling drag
        windmilling_drag_coefficient = windmilling_drag(vehicle,state)

        // Compute asymmetry drag   
        asymmetry_drag_coefficient = asymmetry_drag(state, vehicle, windmilling_drag_coefficient)
           
        // Compute l over d ratio for takeoff condition, NO engine failure
        l_over_d = estimate_2ndseg_lift_drag_ratio(vehicle) 
        
        // Compute L over D ratio for takeoff condition, WITH engine failure
        clv2 = maximum_lift_coefficient / (V2_VS_ratio) **2
        cdv2_all_engine = clv2 / l_over_d
        cdv2 = cdv2_all_engine + asymmetry_drag_coefficient + windmilling_drag_coefficient
        l_over_d_v2 = clv2 / cdv2
    
        // Compute 2nd segment climb gradient
        second_seg_climb_gradient = thrust / (weight*sea_level_gravity) - 1. / l_over_d_v2

        return takeoff_field_length, second_seg_climb_gradient

    else:
        // return only takeoff_field_length
        return takeoff_field_length

After Change

        // Compute 2nd segment climb gradient
        second_seg_climb_gradient = thrust / (weight*sea_level_gravity) - 1. / l_over_d_v2

        return takeoff_field_length[0][0], second_seg_climb_gradient[0][0]

    else:
        // return only takeoff_field_length
        return takeoff_field_length[0][0]

Instances