Python Data.propulsion方法代码示例

# 需要导入模块: from SUAVE.Structure import Data [as 别名]
# 或者: from SUAVE.Structure.Data import propulsion [as 别名]

#.........这里部分代码省略.........
        # Using semi-empirical method for maximum lift coefficient calculation
        from SUAVE.Methods.Aerodynamics.Lift.High_lift_correlations import compute_max_lift_coeff

        # Condition to CLmax calculation: 90KTAS @ 10000ft, ISA
        p_stall , T_stall , rho_stall , a_stall , mew_stall  = atmo.compute_values(10000. * Units.ft)
        conditions = Data()
        conditions.rho = rho_stall
        conditions.mew = mew_stall
        conditions.V = 90. * Units.knots
        try:
            maximum_lift_coefficient, induced_drag_high_lift = compute_max_lift_coeff(config,conditions)
            config.maximum_lift_coefficient = maximum_lift_coefficient
        except:
            raise ValueError, "Maximum lift coefficient calculation error. Please, check inputs"

    # ==============================================
    # Computing speeds (Vs, V2, 0.7*V2)
    # ==============================================
    stall_speed = (2 * weight * sea_level_gravity / (rho * reference_area * maximum_lift_coefficient)) ** 0.5
    V2_speed    = V2_VS_ratio * stall_speed
    speed_for_thrust  = 0.70 * V2_speed

    # ==============================================
    # Determining vehicle number of engines
    # ==============================================
    engine_number = 0.
    for propulsor in vehicle.Propulsors : # may have than one propulsor
        engine_number += propulsor.no_of_engines
    if engine_number == 0:
        raise ValueError, "No engine found in the vehicle"

    # ==============================================
    # Getting engine thrust
    # ==============================================
    #state = Data()
    #state.q  = np.atleast_1d(0.5 * rho * speed_for_thrust**2)
    #state.g0 = np.atleast_1d(sea_level_gravity)
    #state.V  = np.atleast_1d(speed_for_thrust)
    #state.M  = np.atleast_1d(speed_for_thrust/ a_delta_ISA)
    #state.T  = np.atleast_1d(T_delta_ISA)
    #state.p  = np.atleast_1d(p)
    eta      = np.atleast_1d(1.)
    conditions = Data()
    conditions.freestream = Data()
    conditions.propulsion = Data()

    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_delta_ISA)])
    conditions.freestream.temperature = np.array([np.atleast_1d(T_delta_ISA)])
    conditions.freestream.pressure = np.array([np.atleast_1d(p)])
    conditions.propulsion.throttle = np.array([np.atleast_1d(1.)])   

    thrust, mdot, P = vehicle.propulsion_model(eta, conditions) # total thrust

    # ==============================================
    # Calculate takeoff distance
    # ==============================================

    # Defining takeoff distance equations coefficients
    try:
        takeoff_constants = config.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 / 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

    # return
    return takeoff_field_length

# 需要导入模块: from SUAVE.Structure import Data [as 别名]
# 或者: from SUAVE.Structure.Data import propulsion [as 别名]
def main():
    vehicle = SUAVE.Vehicle()# Create the vehicle for testing
    
    # Parameters Required
    vehicle.envelope.ultimate_load                      = 3.5                             # Ultimate load
    vehicle.mass_properties.max_takeoff                 = 79015.8 * Units.kilograms       # Maximum takeoff weight in kilograms
    vehicle.mass_properties.max_zero_fuel               = 79015.8 * 0.9 * Units.kilograms # Maximum zero fuel weight in kilograms
    vehicle.envelope.limit_load                         = 1.5                             # Limit Load
    
    turbofan = SUAVE.Components.Propulsors.TurboFanPASS()
    turbofan.tag = 'Turbo Fan'    
    turbofan.number_of_engines   = 2.                              # Number of engines on the aircraft
    turbofan.thrust.design  = 1000.   * Units.newton    # Define Thrust in Newtons    
    vehicle.append_component(turbofan) 

    vehicle.passengers                                  = 170.                            # Number of passengers
    vehicle.mass_properties.cargo                       = 0.  * Units.kilogram            # Mass of cargo
    vehicle.systems.control                             = "fully powered"                 # Specify fully powered, partially powered or anything else is fully aerodynamic
    vehicle.systems.accessories                         = "medium-range"                  # Specify what type of aircraft you have
    
    vehicle.reference_area                              = 124.862  * Units.meter**2  # Wing gross area in square meters
    wing = SUAVE.Components.Wings.Wing()
    wing.tag = 'Main Wing'
    wing.spans.projected          = 50.      * Units.meter     # Span in meters
    wing.taper                    = 0.2                        # Taper ratio
    wing.thickness_to_chord       = 0.08                       # Thickness-to-chord ratio
    wing.sweep                    = .4363323 * Units.rad       # sweep angle in degrees
    wing.chords.root              = 15.      * Units.meter     # Wing root chord length
    wing.chords.mean_aerodynamic  = 10.      * Units.meters    # Length of the mean aerodynamic chord of the wing
    wing.position                 = [20,0,0] * Units.meters    # Location of main wing from origin of the vehicle
    wing.aerodynamic_center       = [3,0,0]  * Units.meters    # Location of aerodynamic center from origin of the main wing
    vehicle.append_component(wing)
    
    fuselage = SUAVE.Components.Fuselages.Fuselage()
    fuselage.tag = 'Fuselage'    
    fuselage.areas.wetted             = 688.64    * Units.meter**2  # Fuselage wetted area 
    fuselage.differential_pressure    = 55960.5   * Units.pascal    # Maximum differential pressure
    fuselage.width                    = 4.        * Units.meter     # Width of the fuselage
    fuselage.heights.maximum          = 4.        * Units.meter     # Height of the fuselage
    fuselage.lengths.total            = 58.4      * Units.meter     # Length of the fuselage
    fuselage.number_coach_seats       = 200.       
    vehicle.append_component(fuselage)
    
    wing = SUAVE.Components.Wings.Wing()
    wing.tag = 'Horizontal Stabilizer'    
    wing.areas.reference          = 75.     * Units.meters**2 # Area of the horizontal tail
    wing.spans.projected          = 15.     * Units.meters    # Span of the horizontal tail
    wing.sweep                    = 38.     * Units.deg       # Sweep of the horizontal tail
    wing.chords.mean_aerodynamic  = 5.      * Units.meters    # Length of the mean aerodynamic chord of the horizontal tail
    wing.thickness_to_chord       = 0.07                      # Thickness-to-chord ratio of the horizontal tail
    wing.areas.exposed            = 199.7792                  # Exposed area of the horizontal tail
    wing.areas.wetted             = 249.724                   # Wetted area of the horizontal tail
    wing.position                 = [45,0,0]                  # Location of horizontal tail from origin of the vehicle
    wing.aerodynamic_center       = [3,0,0]                   # Location of aerodynamic center from origin of the horizontal tail
    vehicle.append_component(wing)    
    
    wing = SUAVE.Components.Wings.Wing()
    wing.tag = 'Vertical Stabilizer'    
    wing.areas.reference     = 60.     * Units.meters**2 # Area of the vertical tail
    wing.spans.projected     = 15.     * Units.meters    # Span of the vertical tail
    wing.thickness_to_chord  = 0.07                      # Thickness-to-chord ratio of the vertical tail
    wing.sweep               = 40.     * Units.deg       # Sweep of the vertical tail
    wing.t_tail              = "false"                   # Set to "yes" for a T-tail
    vehicle.append_component(wing)   
    
    weight = Tube_Wing.empty(vehicle)
    
    actual = Data()
    actual.payload = 17349.9081525
    actual.pax = 15036.5870655
    actual.bag = 2313.321087
    actual.fuel = -6993.89102491
    actual.empty = 68659.7828724
    actual.wing = 27694.192985
    actual.fuselage = 11504.5186408
    actual.propulsion = 88.3696093424
    actual.landing_gear = 3160.632
    actual.systems = 16655.7076511
    actual.wt_furnish = 7466.1304102
    actual.horizontal_tail = 2191.30720639
    actual.vertical_tail = 5260.75341411
    actual.rudder = 2104.30136565    
    
    error = Data()
    error.payload = (actual.payload - weight.payload)/actual.payload
    error.pax = (actual.pax - weight.pax)/actual.pax
    error.bag = (actual.bag - weight.bag)/actual.bag
    error.fuel = (actual.fuel - weight.fuel)/actual.fuel
    error.empty = (actual.empty - weight.empty)/actual.empty
    error.wing = (actual.wing - weight.wing)/actual.wing
    error.fuselage = (actual.fuselage - weight.fuselage)/actual.fuselage
    error.propulsion = (actual.propulsion - weight.propulsion)/actual.propulsion
    error.landing_gear = (actual.landing_gear - weight.landing_gear)/actual.landing_gear
    error.systems = (actual.systems - weight.systems)/actual.systems
    error.wt_furnish = (actual.wt_furnish - weight.wt_furnish)/actual.wt_furnish
    error.horizontal_tail = (actual.horizontal_tail - weight.horizontal_tail)/actual.horizontal_tail
    error.vertical_tail = (actual.vertical_tail - weight.vertical_tail)/actual.vertical_tail
    error.rudder = (actual.rudder - weight.rudder)/actual.rudder
      
    for k,v in error.items():
#.........这里部分代码省略.........

# 需要导入模块: from SUAVE.Structure import Data [as 别名]
# 或者: from SUAVE.Structure.Data import propulsion [as 别名]
    def __defaults__(self):
        self.tag = 'Aerodynamic Segment'

        # atmosphere and planet
        self.planet     = None
        self.atmosphere = None
        self.start_time = time.gmtime()


        # --- Conditions and Unknowns

        # user shouldn't change these in an input script
        # only used for processing / post processing
        # they will be shared with analysis modules and meaningful naming is important

        # base matricies
        # use a trivial operation to copy the array
        ones_1col = np.ones([1,1])
        ones_2col = np.ones([1,2])
        ones_3col = np.ones([1,3])


        # --- Conditions

        # setup conditions
        conditions = Data()
        conditions.frames       = Data()
        conditions.freestream   = Data()
        conditions.aerodynamics = Data()
        conditions.propulsion   = Data()
        conditions.weights      = Data()
        conditions.energies     = Data()
        self.conditions = conditions

        # inertial frame conditions
        conditions.frames.inertial = Data()
        conditions.frames.inertial.position_vector      = ones_3col * 0
        conditions.frames.inertial.velocity_vector      = ones_3col * 0
        conditions.frames.inertial.acceleration_vector  = ones_3col * 0
        conditions.frames.inertial.gravity_force_vector = ones_3col * 0
        conditions.frames.inertial.total_force_vector   = ones_3col * 0
        conditions.frames.inertial.time                 = ones_1col * 0

        # wind frame conditions
        conditions.frames.wind = Data()
        conditions.frames.wind.body_rotations           = ones_3col * 0   # rotations in [X,Y,Z] -> [phi,theta,psi]
        conditions.frames.wind.velocity_vector          = ones_3col * 0
        conditions.frames.wind.lift_force_vector        = ones_3col * 0
        conditions.frames.wind.drag_force_vector        = ones_3col * 0
        conditions.frames.wind.transform_to_inertial    = np.empty([0,0,0])

        # body frame conditions
        conditions.frames.body = Data()
        conditions.frames.body.inertial_rotations       = ones_3col * 0    # rotations in [X,Y,Z] -> [phi,theta,psi]
        conditions.frames.body.thrust_force_vector      = ones_3col * 0
        conditions.frames.body.transform_to_inertial    = np.empty([0,0,0])

        # planet frame conditions
        conditions.frames.planet = Data()
        conditions.frames.planet.start_time      = None
        conditions.frames.planet.latitude        = ones_1col * 0
        conditions.frames.planet.longitude       = ones_1col * 0

        # freestream conditions
        conditions.freestream.velocity           = ones_1col * 0
        conditions.freestream.mach_number        = ones_1col * 0
        conditions.freestream.pressure           = ones_1col * 0
        conditions.freestream.temperature        = ones_1col * 0
        conditions.freestream.density            = ones_1col * 0
        conditions.freestream.speed_of_sound     = ones_1col * 0
        conditions.freestream.viscosity          = ones_1col * 0
        conditions.freestream.altitude           = ones_1col * 0
        conditions.freestream.gravity            = ones_1col * 0
        conditions.freestream.reynolds_number    = ones_1col * 0
        conditions.freestream.dynamic_pressure   = ones_1col * 0

        # aerodynamics conditions
        conditions.aerodynamics.angle_of_attack  = ones_1col * 0
        conditions.aerodynamics.side_slip_angle  = ones_1col * 0
        conditions.aerodynamics.roll_angle       = ones_1col * 0
        conditions.aerodynamics.lift_coefficient = ones_1col * 0
        conditions.aerodynamics.drag_coefficient = ones_1col * 0
        conditions.aerodynamics.lift_breakdown   = Data()
        conditions.aerodynamics.drag_breakdown   = Data()

        # propulsion conditions
        conditions.propulsion.throttle           = ones_1col * 0
        conditions.propulsion.fuel_mass_rate     = ones_1col * 0
        conditions.propulsion.battery_energy     = ones_1col * 0
        conditions.propulsion.thrust_breakdown   = Data()

        # weights conditions
        conditions.weights.total_mass            = ones_1col * 0
        conditions.weights.weight_breakdown      = Data()

        # energy conditions
        conditions.energies.total_energy         = ones_1col * 0
        conditions.energies.total_efficiency     = ones_1col * 0
        conditions.energies.gravity_energy       = ones_1col * 0
        conditions.energies.propulsion_power     = ones_1col * 0
#.........这里部分代码省略.........

# 需要导入模块: from SUAVE.Structure import Data [as 别名]
# 或者: from SUAVE.Structure.Data import propulsion [as 别名]
def main():

    # ------------------------------------------------------------------
    #   Propulsor
    # ------------------------------------------------------------------
    
    # build network
    net = Solar_Network()
    net.number_motors = 1.
    net.nacelle_dia   = 0.2
    
    # Component 1 the Sun?
    sun = SUAVE.Components.Energy.Processes.Solar_Radiation()
    net.solar_flux = sun
    
    # Component 2 the solar panels
    panel = SUAVE.Components.Energy.Converters.Solar_Panel()
    panel.area                 = 100 * Units.m
    panel.efficiency           = 0.18
    panel.mass_properties.mass = panel.area*.600
    net.solar_panel            = panel
    
    # Component 3 the ESC
    esc = SUAVE.Components.Energy.Distributors.Electronic_Speed_Controller()
    esc.efficiency = 0.95 # Gundlach for brushless motors
    net.esc       = esc
    
    # Component 5 the Propeller
    
    #Propeller design specs
    design_altitude = 0.0 * Units.km
    Velocity        = 10.0  # freestream m/s
    RPM             = 5887
    Blades          = 2.0
    Radius          = .4064
    Hub_Radius      = 0.05
    Design_Cl       = 0.7
    Thrust          = 0.0 #Specify either thrust or power to design for
    Power           = 7500.  #Specify either thrust or power to design for
    
    # Design the Propeller
    prop_attributes = Data()
    prop_attributes.number_blades       = Blades 
    prop_attributes.freestream_velocity = Velocity
    prop_attributes.angular_velocity    = RPM*(2.*np.pi/60.0)
    prop_attributes.tip_radius          = Radius
    prop_attributes.hub_radius          = Hub_Radius
    prop_attributes.design_Cl           = Design_Cl 
    prop_attributes.design_altitude     = design_altitude
    prop_attributes.design_thrust       = Thrust
    prop_attributes.design_power        = Power
    prop_attributes                     = propeller_design(prop_attributes)
    
    # Create and attach this propeller
    prop                 = SUAVE.Components.Energy.Converters.Propeller()
    prop.prop_attributes = prop_attributes
    net.propeller        = prop
    
    # Component 4 the Motor
    motor = SUAVE.Components.Energy.Converters.Motor()
    motor.resistance           = 0.01
    motor.no_load_current      = 8.0
    motor.speed_constant       = 140.*(2.*np.pi/60.) # RPM/volt converted to rad/s     
    motor.propeller_radius     = prop.prop_attributes.tip_radius
    motor.propeller_Cp         = prop.prop_attributes.Cp
    motor.gear_ratio           = 1.
    motor.gearbox_efficiency   = 1.
    motor.expected_current     = 260.
    motor.mass_properties.mass = 2.0
    net.motor                  = motor   
    
    # Component 6 the Payload
    payload = SUAVE.Components.Energy.Peripherals.Payload()
    payload.power_draw           = 0. #Watts 
    payload.mass_properties.mass = 0. * Units.kg
    net.payload                  = payload
    
    # Component 7 the Avionics
    avionics = SUAVE.Components.Energy.Peripherals.Avionics()
    avionics.power_draw = 0. #Watts  
    net.avionics        = avionics      
    
    # Component 8 the Battery
    bat = SUAVE.Components.Energy.Storages.Battery()
    bat.mass_properties.mass = 50.  #kg
    bat.type = 'Li-Ion'
    bat.resistance = 0.0
    net.battery = bat
    
    #Component 9 the system logic controller and MPPT
    logic = SUAVE.Components.Energy.Distributors.Solar_Logic()
    logic.system_voltage  = 50.0
    logic.MPPT_efficiency = 0.95
    net.solar_logic       = logic
    
    # Setup the conditions to run the network
    conditions                 = Data()
    conditions.propulsion      = Data()
    conditions.freestream      = Data()
    conditions.frames          = Data()
#.........这里部分代码省略.........

# 需要导入模块: from SUAVE.Structure import Data [as 别名]
# 或者: from SUAVE.Structure.Data import propulsion [as 别名]
def main():
    
    # This script could fail if either the design or analysis scripts fail,
    # in case of failure check both. The design and analysis powers will 
    # differ because of karman-tsien compressibility corrections in the 
    # analysis scripts
    
    # Design the Propeller
    prop_attributes = Data()
    prop_attributes.number_blades       = 2.0 
    prop_attributes.freestream_velocity = 50.0
    prop_attributes.angular_velocity    = 2000.*(2.*np.pi/60.0)
    prop_attributes.tip_radius          = 1.5
    prop_attributes.hub_radius          = 0.05
    prop_attributes.design_Cl           = 0.7 
    prop_attributes.design_altitude     = 0.0 * Units.km
    prop_attributes.design_thrust       = 0.0
    prop_attributes.design_power        = 7000.
    prop_attributes                     = propeller_design(prop_attributes)    

    # Find the operating conditions
    atmosphere = SUAVE.Attributes.Atmospheres.Earth.US_Standard_1976()
    p, T, rho, a, mu = atmosphere.compute_values(prop_attributes.design_altitude)
    
    V = prop_attributes.freestream_velocity
    
    conditions = Data()
    conditions.freestream = Data()
    conditions.propulsion = Data()
    conditions.freestream.density        = np.array([rho])
    conditions.freestream.viscosity      = np.array([mu])
    conditions.freestream.velocity       = np.array([[V]])
    conditions.freestream.speed_of_sound = np.array([a])
    conditions.freestream.temperature    = np.array([T])
    conditions.propulsion.throttle       = np.array([[1.0]])
    
    # Create and attach this propeller
    prop                 = SUAVE.Components.Energy.Converters.Propeller()
    prop.prop_attributes = prop_attributes    
    prop.inputs.omega    = prop_attributes.angular_velocity
    
    F, Q, P, Cplast = prop.spin(conditions)
    
    # Truth values
    F_truth      = 166.41590262
    Q_truth      = 45.21732911
    P_truth      = 9470.2952633 # Over 9000!
    Cplast_truth = 0.00085898
    
    error = Data()
    error.Thrust  = np.max(np.abs(F-F_truth))
    error.Power   = np.max(np.abs(P-P_truth))
    error.Torque  = np.max(np.abs(Q-Q_truth))
    error.Cp      = np.max(np.abs(Cplast-Cplast_truth))   
    
    print  error
    
    for k,v in error.items():
        assert(np.abs(v)<0.001)
     
    return