Python Data.numerics方法代码示例

# 需要导入模块: from SUAVE.Core import Data [as 别名]
# 或者: from SUAVE.Core.Data import numerics [as 别名]
def compute_ducted_fan_geometry(ducted_fan, conditions):
    """ SUAVE.Methods.Geometry.Two_Dimensional.Planform.wing_fuel_volume(wing):
        Estimates wing fuel capacity based in correlation methods.



    """

    # unpack
    thrust            = ducted_fan.thrust
    fan_nozzle        = ducted_fan.fan_nozzle
    mass_flow         = thrust.mass_flow_rate_design

    #evaluate engine at these conditions
    state=Data()
    state.conditions=conditions
    state.numerics= Data()
    ducted_fan.evaluate_thrust(state)
    
    #determine geometry
    U0       = conditions.freestream.velocity
    rho0     = conditions.freestream.density
    Ue       = fan_nozzle.outputs.velocity
    rhoe     = fan_nozzle.outputs.density
    Ae       = mass_flow[0][0]/(rhoe[0][0]*Ue[0][0]) #ducted fan nozzle exit area
    A0       = (mass_flow/(rho0*U0))[0][0]
    

   
    ducted_fan.areas.maximum = 1.2*Ae/fan_nozzle.outputs.area_ratio[0][0]
    ducted_fan.nacelle_diameter = 2.1*((ducted_fan.areas.maximum/np.pi)**.5)

    ducted_fan.engine_length    = 1.5*ducted_fan.nacelle_diameter
    ducted_fan.areas.wetted     = ducted_fan.nacelle_diameter*ducted_fan.engine_length*np.pi

# 需要导入模块: from SUAVE.Core import Data [as 别名]
# 或者: from SUAVE.Core.Data import numerics [as 别名]
def energy_network():
    
    # ------------------------------------------------------------------
    #   Evaluation Conditions
    # ------------------------------------------------------------------      
    
    # setup conditions
    conditions = SUAVE.Analyses.Mission.Segments.Conditions.Aerodynamics()   
    planet     = SUAVE.Attributes.Planets.Earth()
    ones_1col             = np.ones([1,1])   
    conditions.freestream = Data()
    conditions.propulsion = Data()
    
    
    # vacuum conditions
    vac                   = conditions.freestream
    vac.altitude          = ones_1col*0.0
    vac.gravity           = ones_1col*planet.sea_level_gravity
    vac.pressure          = ones_1col*0.0
        
    # propulsion conditions
    conditions.propulsion.throttle  =  ones_1col*1.0

    # ------------------------------------------------------------------
    #   Design/sizing conditions
    # ------------------------------------------------------------------    
    
    # setup conditions
    conditions_sls = SUAVE.Analyses.Mission.Segments.Conditions.Aerodynamics()      
    ones_1col                         = np.ones([1,1]) 
    conditions_sls.freestream         = Data()
    conditions_sls.propulsion         = Data()
    
    # freestream conditions
    SLS                               = conditions_sls.freestream
    SLS.altitude                      = ones_1col*0.0
    
    atmosphere                        = SUAVE.Analyses.Atmospheric.US_Standard_1976()
    atmo_data                         = atmosphere.compute_values(SLS.altitude,0,True) 
    
    SLS.pressure                      = ones_1col*atmo_data.pressure
    SLS.temperature                   = ones_1col*atmo_data.temperature
    SLS.density                       = ones_1col*atmo_data.density
    SLS.speed_of_sound                = ones_1col*atmo_data.speed_of_sound
    SLS.gravity                       = ones_1col*planet.sea_level_gravity
    
    # propulsion conditions
    conditions_sls.propulsion.throttle = ones_1col*1.0

    # setting states
    state_sls               = Data()
    state_sls.numerics      = Data()
    state_sls.conditions    = conditions_sls
    state_vacuum            = Data()
    state_vacuum.numerics   = Data()
    state_vacuum.conditions = conditions

    # ------------------------------------------------------------------
    #  F-1 Liquid Rocket Network
    # ------------------------------------------------------------------    
    
    # instantiate the ramjet network
    liquid_rocket = SUAVE.Components.Energy.Networks.Liquid_Rocket()
    liquid_rocket.tag = 'liquid_rocket'
    
    # setup
    liquid_rocket.number_of_engines = 1.0
    liquid_rocket.area_throat       = 0.6722
    liquid_rocket.contraction_ratio = 2.8956
    liquid_rocket.expansion_ratio   = 16.0
    
    # ------------------------------------------------------------------
    #   Component 1 - Combustor
     
    # instantiate    
    combustor = SUAVE.Components.Energy.Converters.Rocket_Combustor()   
    combustor.tag = 'combustor'
    
    # setup  
    combustor.propellant_data                = SUAVE.Attributes.Propellants.LOX_RP1()
    combustor.inputs.combustion_pressure     = 7000000.0 
    
    # add to network
    liquid_rocket.append(combustor)

    # ------------------------------------------------------------------
    #  Component 2 - Core Nozzle
    
    # instantiate
    nozzle = SUAVE.Components.Energy.Converters.de_Laval_Nozzle()   
    nozzle.tag = 'core_nozzle'
    
    # setup
    nozzle.polytropic_efficiency = 1.0
    nozzle.expansion_ratio       = liquid_rocket.expansion_ratio
    nozzle.area_throat           = liquid_rocket.area_throat
    nozzle.pressure_ratio        = 1.0
    
    
    # add to network
#.........这里部分代码省略.........

# 需要导入模块: from SUAVE.Core import Data [as 别名]
# 或者: from SUAVE.Core.Data import numerics [as 别名]
def energy_network():
    
    # ------------------------------------------------------------------
    #   Evaluation Conditions
    # ------------------------------------------------------------------    
    
    # Conditions        
    ones_1col = np.ones([1,1])       
    alt                                                = 10.0
    
    # Setup conditions
    planet     = SUAVE.Attributes.Planets.Earth()   
    atmosphere                       = SUAVE.Analyses.Atmospheric.US_Standard_1976()
    atmo_data                        = atmosphere.compute_values(alt,0,True) 
    working_fluid                    = SUAVE.Attributes.Gases.Air()    
    conditions = SUAVE.Analyses.Mission.Segments.Conditions.Aerodynamics()
  
    # freestream conditions
    conditions.freestream.altitude                     = ones_1col*alt   
    conditions.freestream.mach_number                  = ones_1col*0.8
    conditions.freestream.pressure                     = ones_1col*atmo_data.pressure
    conditions.freestream.temperature                  = ones_1col*atmo_data.temperature
    conditions.freestream.density                      = ones_1col*atmo_data.density
    conditions.freestream.dynamic_viscosity            = ones_1col*atmo_data.dynamic_viscosity
    conditions.freestream.gravity                      = ones_1col*planet.compute_gravity(alt)
    conditions.freestream.isentropic_expansion_factor  = ones_1col*working_fluid.compute_gamma(atmo_data.temperature,atmo_data.pressure)                                                                                             
    conditions.freestream.Cp                           = ones_1col*working_fluid.compute_cp(atmo_data.temperature,atmo_data.pressure)
    conditions.freestream.R                            = ones_1col*working_fluid.gas_specific_constant
    conditions.freestream.speed_of_sound               = ones_1col*atmo_data.speed_of_sound
    conditions.freestream.velocity                     = conditions.freestream.mach_number*conditions.freestream.speed_of_sound
    conditions.velocity                                = conditions.freestream.mach_number*conditions.freestream.speed_of_sound
    conditions.q                                       = 0.5*conditions.freestream.density*conditions.velocity**2
    conditions.g0                                      = conditions.freestream.gravity
    
    # propulsion conditions
    conditions.propulsion.throttle                     =  ones_1col*1.0
        
    # ------------------------------------------------------------------
    #   Design/sizing conditions
    # ------------------------------------------------------------------    
    
    # Conditions        
    ones_1col = np.ones([1,1])    
    alt_size  = 10000.0
    # Setup conditions
    planet     = SUAVE.Attributes.Planets.Earth()   
    atmosphere                       = SUAVE.Analyses.Atmospheric.US_Standard_1976()
    atmo_data                        = atmosphere.compute_values(alt_size,0,True) 
    working_fluid                    = SUAVE.Attributes.Gases.Air()    
    conditions_sizing = SUAVE.Analyses.Mission.Segments.Conditions.Aerodynamics()

    # freestream conditions
    conditions_sizing.freestream.altitude                     = ones_1col*alt_size     
    conditions_sizing.freestream.mach_number                  = ones_1col*0.8
    conditions_sizing.freestream.pressure                     = ones_1col*atmo_data.pressure
    conditions_sizing.freestream.temperature                  = ones_1col*atmo_data.temperature
    conditions_sizing.freestream.density                      = ones_1col*atmo_data.density
    conditions_sizing.freestream.dynamic_viscosity            = ones_1col*atmo_data.dynamic_viscosity
    conditions_sizing.freestream.gravity                      = ones_1col*planet.compute_gravity(alt_size)
    conditions_sizing.freestream.isentropic_expansion_factor  = ones_1col*working_fluid.compute_gamma(atmo_data.temperature,atmo_data.pressure)                                                                                             
    conditions_sizing.freestream.Cp                           = ones_1col*working_fluid.compute_cp(atmo_data.temperature,atmo_data.pressure)
    conditions_sizing.freestream.R                            = ones_1col*working_fluid.gas_specific_constant
    conditions_sizing.freestream.speed_of_sound               = ones_1col*atmo_data.speed_of_sound
    conditions_sizing.freestream.velocity                     = conditions_sizing.freestream.mach_number*conditions_sizing.freestream.speed_of_sound
    conditions_sizing.velocity                                = conditions_sizing.freestream.mach_number*conditions_sizing.freestream.speed_of_sound
    conditions_sizing.q                                       = 0.5*conditions_sizing.freestream.density*conditions_sizing.velocity**2
    conditions_sizing.g0                                      = conditions_sizing.freestream.gravity
    
    # propulsion conditions
    conditions_sizing.propulsion.throttle                    =  ones_1col*1.0

    state_sizing                = Data()
    state_sizing.numerics       = Data()
    state_sizing.conditions     = conditions_sizing
    state_off_design            = Data()
    state_off_design.numerics   = Data()
    state_off_design.conditions = conditions

    # ------------------------------------------------------------------
    #   Turbofan Network
    # ------------------------------------------------------------------    
    
    # Instantiate the gas turbine network
    turbofan     = SUAVE.Components.Energy.Networks.Turbofan()
    turbofan.tag = 'turbofan'
    
    # setup
    turbofan.bypass_ratio      = 5.4
    turbofan.number_of_engines = 2.0
    turbofan.engine_length     = 2.5
    turbofan.nacelle_diameter  = 1.580
    
    # working fluid
    turbofan.working_fluid = SUAVE.Attributes.Gases.Air()
    
    
    # ------------------------------------------------------------------
    #   Component 1 - Ram
        
    # instantiate
#.........这里部分代码省略.........

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

#.........这里部分代码省略.........
    # ==============================================
    # Determining vehicle maximum lift coefficient
    # ==============================================
    try:   # aircraft maximum lift informed by user
        maximum_lift_coefficient = vehicle.maximum_lift_coefficient
    except:
        # Using semi-empirical method for maximum lift coefficient calculation
        from SUAVE.Methods.Aerodynamics.Fidelity_Zero.Lift import compute_max_lift_coeff

        # Condition to CLmax calculation: 90KTAS @ 10000ft, ISA
        conditions  = atmo.compute_values(10000. * Units.ft)
        conditions.freestream=Data()
        conditions.freestream.density   = conditions.density
        conditions.freestream.dynamic_viscosity = conditions.dynamic_viscosity
        conditions.freestream.velocity  = 90. * Units.knots
        try:
            maximum_lift_coefficient, induced_drag_high_lift = compute_max_lift_coeff(vehicle,conditions)
            vehicle.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.number_of_engines
    if engine_number == 0:
        raise ValueError, "No engine found in the vehicle"

    # ==============================================
    # Getting engine thrust
    # ==============================================
    
    state = Data()
    state.conditions = conditions
    state.numerics   = 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.))
    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
        p
    takeoff_field_length = takeoff_field_length * Units.ft
    # return
    return takeoff_field_length

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

#.........这里部分代码省略.........
    #flow through the low pressure turbine
    low_pressure_turbine(conditions)
    
    #link the core nozzle to the low pressure turbine
    core_nozzle.inputs.stagnation_temperature              = low_pressure_turbine.outputs.stagnation_temperature
    core_nozzle.inputs.stagnation_pressure                 = low_pressure_turbine.outputs.stagnation_pressure
    
    #flow through the core nozzle
    core_nozzle(conditions)
   
    #link the fan nozzle to the fan
    fan_nozzle.inputs.stagnation_temperature               = fan.outputs.stagnation_temperature
    fan_nozzle.inputs.stagnation_pressure                  = fan.outputs.stagnation_pressure
    
    #flow through the fan nozzle
    fan_nozzle(conditions)
    
    # compute the thrust using the thrust component
    #link the thrust component to the fan nozzle
    thrust.inputs.fan_exit_velocity                        = fan_nozzle.outputs.velocity
    thrust.inputs.fan_area_ratio                           = fan_nozzle.outputs.area_ratio
    thrust.inputs.fan_nozzle                               = fan_nozzle.outputs
    
    #link the thrust component to the core nozzle
    thrust.inputs.core_exit_velocity                       = core_nozzle.outputs.velocity
    thrust.inputs.core_area_ratio                          = core_nozzle.outputs.area_ratio
    thrust.inputs.core_nozzle                              = core_nozzle.outputs
    
    #link the thrust component to the combustor
    thrust.inputs.fuel_to_air_ratio                        = combustor.outputs.fuel_to_air_ratio
    
    #link the thrust component to the low pressure compressor 
    thrust.inputs.total_temperature_reference              = low_pressure_compressor.outputs.stagnation_temperature
    thrust.inputs.total_pressure_reference                 = low_pressure_compressor.outputs.stagnation_pressure
    thrust.inputs.number_of_engines                        = number_of_engines
    thrust.inputs.bypass_ratio                             = bypass_ratio
    thrust.inputs.flow_through_core                        =  1./(1.+bypass_ratio) #scaled constant to turn on core thrust computation
    thrust.inputs.flow_through_fan                         =  bypass_ratio/(1.+bypass_ratio) #scaled constant to turn on fan thrust computation     

    #compute the thrust
    thrust.size(conditions)
    
    #determine geometry; 
    mass_flow         = thrust.mass_flow_rate_design
    mass_flow_fan     = mass_flow*bypass_ratio
    
 
 
    U0                = conditions.freestream.velocity
    gamma             = ram.outputs.isentropic_expansion_factor
    R                 = ram.outputs.universal_gas_constant
    rho0              = conditions.freestream.density
    
    rho5_fan          = fan_nozzle.outputs.density
    U5_fan            = fan_nozzle.outputs.velocity
    
    rho5_core         = core_nozzle.outputs.density
    U5_core           = core_nozzle.outputs.velocity
    
    
    
  
    
    #update the design thrust value
    turbofan.design_thrust = thrust.total_design
    
    #compute the sls_thrust
    #call the atmospheric model to get the conditions at the specified altitude
    atmosphere_sls = SUAVE.Analyses.Atmospheric.US_Standard_1976()
    p,T,rho,a,mu = atmosphere_sls.compute_values(0.0,0.0)

    # setup conditions
    conditions_sls = SUAVE.Analyses.Mission.Segments.Conditions.Aerodynamics()            

    # freestream conditions    
    conditions_sls.freestream.altitude           = np.atleast_1d(0.)
    conditions_sls.freestream.mach_number        = np.atleast_1d(0.01)
    conditions_sls.freestream.pressure           = np.atleast_1d(p)
    conditions_sls.freestream.temperature        = np.atleast_1d(T)
    conditions_sls.freestream.density            = np.atleast_1d(rho)
    conditions_sls.freestream.dynamic_viscosity  = np.atleast_1d(mu)
    conditions_sls.freestream.gravity            = np.atleast_1d(9.81)
    conditions_sls.freestream.gamma              = np.atleast_1d(1.4)
    conditions_sls.freestream.Cp                 = 1.4*(p/(rho*T))/(1.4-1)
    conditions_sls.freestream.R                  = p/(rho*T)
    conditions_sls.freestream.speed_of_sound     = np.atleast_1d(a)
    conditions_sls.freestream.velocity           = np.atleast_1d(a*0.01)
    
    # propulsion conditions
    conditions_sls.propulsion.throttle           =  np.atleast_1d(1.0)    
    
    #size the turbofan

    state_sls = Data()
    state_sls.numerics = Data()
    state_sls.conditions = conditions_sls   
    results_sls = turbofan.evaluate_thrust(state_sls)
    
    
    turbofan.sealevel_static_thrust = results_sls.thrust_force_vector[0,0] / number_of_engines

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

#.........这里部分代码省略.........
        # Condition to CLmax calculation: 90KTAS @ 10000ft, ISA
        conditions  = atmo.compute_values(10000. * Units.ft)
        conditions.freestream=Data()
        conditions.freestream.density   = conditions.density
        conditions.freestream.dynamic_viscosity = conditions.dynamic_viscosity
        conditions.freestream.velocity  = 90. * Units.knots
        try:
            maximum_lift_coefficient, induced_drag_high_lift = compute_max_lift_coeff(vehicle,conditions)
            vehicle.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.number_of_engines
    if engine_number == 0:
        raise ValueError, "No engine found in the vehicle"

    # ==============================================
    # 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.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:

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

#.........这里部分代码省略.........
    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.Batteries.Constant_Mass.Lithium_Ion()
    batterymass = 50.  #kg
    bat.type = 'Li-Ion'
    bat.resistance = 0.0
    bat.energy_density = 250.
    initialize_from_mass(bat,batterymass)
    bat.current_energy = bat.max_energy
    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
    state            = Data()
    state.conditions = SUAVE.Analyses.Mission.Segments.Conditions.Aerodynamics()
    state.numerics   = SUAVE.Analyses.Mission.Segments.Conditions.Numerics()
    
    conditions = state.conditions
    numerics   = state.numerics
    
    # Calculate atmospheric properties
    atmosphere = SUAVE.Analyses.Atmospheric.US_Standard_1976()
    atmosphere_conditions =  atmosphere.compute_values(prop_attributes.design_altitude)
    
    rho = atmosphere_conditions.density[0,:]
    a   = atmosphere_conditions.speed_of_sound[0,:]
    mu  = atmosphere_conditions.dynamic_viscosity[0,:]
    T   = atmosphere_conditions.temperature[0,:]

    conditions.propulsion.throttle            = np.array([[1.0],[1.0]])
    conditions.freestream.velocity            = np.array([[1.0],[1.0]])
    conditions.freestream.density             = np.array([rho,rho])
    conditions.freestream.dynamic_viscosity   = np.array([mu, mu])
    conditions.freestream.speed_of_sound      = np.array([a, a])
    conditions.freestream.altitude            = np.array([[design_altitude], [design_altitude]])
    conditions.propulsion.battery_energy      = bat.max_energy*np.ones_like(conditions.freestream.altitude)
    conditions.frames.body.inertial_rotations = np.zeros([2,3])
    conditions.frames.inertial.time           = np.array([[0.0],[1.0]])
    numerics.time.integrate                   = np.array([[0, 0],[0, 1]])
    numerics.time.differentiate               = np.array([[0, 0],[0, 1]])
    conditions.frames.planet.start_time       = time.strptime("Sat, Jun 21 06:00:00  2014", "%a, %b %d %H:%M:%S %Y",) 
    conditions.frames.planet.latitude         = np.array([[0.0],[0.0]])
    conditions.frames.planet.longitude        = np.array([[0.0],[0.0]])
    conditions.freestream.temperature         = np.array([T, T])
    conditions.frames.body.transform_to_inertial = np.array([[[ 1.,  0.,  0.],
                                                              [ 0.,  1.,  0.],
                                                              [ 0.,  0.,  1.]],
                                                             [[ 1.,  0.,  0.],
                                                              [ 0.,  1.,  0.],
                                                              [ 0.,  0.,  1.]]])
    conditions.propulsion.propeller_power_coefficient = np.array([[1.], [1.]]) * prop.prop_attributes.Cp
    
    # Run the network and print the results
    results = net(state)
    F       = results.thrust_force_vector
    
    # Truth results
    truth_F   = [[ 545.35952329,  545.35952329]]
    truth_i   = [[ 249.31622624], [ 249.31622624]]
    truth_rpm = [[ 6668.4094191], [ 6668.4094191]]
    truth_bat = [[ 36000000.   ], [ 35987534.18868808]]
    
    error = Data()
    error.Thrust = np.max(np.abs(F[:,0]-truth_F))
    error.RPM = np.max(np.abs(conditions.propulsion.rpm-truth_rpm))
    error.Current  = np.max(np.abs(conditions.propulsion.current-truth_i))
    error.Battery = np.max(np.abs(bat.current_energy-truth_bat))
    
    print(error)
    
    for k,v in list(error.items()):
        assert(np.abs(v)<1e-6)
        
    
    return

# 需要导入模块: from SUAVE.Core import Data [as 别名]
# 或者: from SUAVE.Core.Data import numerics [as 别名]
def print_engine_data(vehicle, filename='engine_data.dat', units="imperial"):
    """This creates a file showing engine information.

    Assumptions:
    One propulsor (can be multiple engines) with 'turbofan' tag.

    Source:
    N/A

    Inputs:
    vehicle.
      tag
      turbofan()              Function to compute thrust and fuel burn rate
      propulsors.turbofan.
        design_thrust         [N]
        engine_length         [m]
        nacelle_diameter      [m]
        thrust.bypass_ratio   [-]
    filename (optional)       <string> Determines the name of the saved file
    units (optional)          <string> Determines the type of units used in the output, options are imperial and si

    Outputs:
    filename                  Saved file with name as above

    Properties Used:
    N/A
    """       
    imperial = False
    SI = False

    if units.lower() == "imperial":
        imperial = True
    elif units.lower() == "si":
        SI = True
    else:
        print "Incorrect system of units selected - choose 'imperial' or 'SI'"
        return

    d_isa_vec = [0, 10]
    if imperial:
        speed_vec = np.linspace(0, 450, 10) * Units.knots
        hp_vec = np.linspace(0, 50000, 11) * Units.ft

    elif SI:
        speed_vec = np.linspace(0, 225, 10) * Units['m/s']
        hp_vec = np.linspace(0, 15000, 11) * Units.m

    speed_vec[0] = 1.

    thrust = np.zeros_like(speed_vec)
    mdot = np.zeros_like(speed_vec)

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


    engine_tag = vehicle.propulsors.turbofan.tag
    design_thrust = vehicle.propulsors.turbofan.design_thrust
    engine_length = vehicle.propulsors.turbofan.engine_length
    nacelle_diameter = vehicle.propulsors.turbofan.nacelle_diameter
    bypass_ratio = vehicle.propulsors.turbofan.thrust.bypass_ratio

    # Considering planet and atmosphere of 1st mission segment
    sea_level_gravity = 9.81  # mission.segments[0].planet.sea_level_gravity
    atmo = SUAVE.Analyses.Atmospheric.US_Standard_1976()  # mission.segments[0].atmosphere

    atmo_values     = atmo.compute_values(0.,0.)
    
    p0   = atmo_values.pressure
    T0   = atmo_values.temperature
    rho0 = atmo_values.density
    a0   = atmo_values.speed_of_sound
    mu0  = atmo_values.dynamic_viscosity    

    state = Data()
    state.conditions = SUAVE.Analyses.Mission.Segments.Conditions.Aerodynamics()
    state.numerics = SUAVE.Analyses.Mission.Segments.Conditions.Numerics()

    # write header of file
    fid = open(filename, 'w')  # Open output file
    fid.write('Output file with engine data \n\n')
    fid.write('  VEHICLE TAG : ' + vehicle.tag + '\n\n')

    for d_isa in d_isa_vec:
        fid.write('\n DELTA ISA: {:3.0f} degC'.format(d_isa) + '\n')

        if imperial:
            fid.write('\n' + 8 * ' ' + '|' + 40 * ' ' + ' SPEED [KTAS]' + 48 * ' ' + '|' + 27 * ' ' + ' SPEED [KTAS]' + 33 * ' ')
            fid.write('\n' + 8 * ' ' + '|')
            fid.write(np.transpose(map('{:10.0f}'.format, speed_vec / Units.knots)))
            fid.write(' |' + 3 * ' ')
            fid.write(np.transpose(map('{:6.0f} '.format, speed_vec / Units.knots)))
            fid.write( '\n HP[ft] |' + 40 * ' ' + ' THRUST [lbf]' + 48 * ' ' + '|' + 30 * ' ' + ' SFC [adm]' + 33 * ' ' + '|\n')

        elif SI:
            fid.write('\n' + 8 * ' ' + '|' + 40 * ' ' + ' SPEED [m/s]' + 48 * ' ' + '|' + 27 * ' ' + ' SPEED [m/s]' + 33 * ' ')
#.........这里部分代码省略.........

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

#.........这里部分代码省略.........
    high_pressure_turbine.inputs.stagnation_pressure       = combustor.outputs.stagnation_pressure
    high_pressure_turbine.inputs.fuel_to_air_ratio         = combustor.outputs.fuel_to_air_ratio
    
    #link the high pressuer turbine to the high pressure compressor
    high_pressure_turbine.inputs.compressor                = high_pressure_compressor.outputs

    #flow through the high pressure turbine
    high_pressure_turbine.inputs.bypass_ratio = 0.0
    high_pressure_turbine.inputs.fan = Data()
    high_pressure_turbine.inputs.fan.work_done = 0.0
    high_pressure_turbine(conditions)
            
    #link the low pressure turbine to the high pressure turbine
    low_pressure_turbine.inputs.stagnation_temperature     = high_pressure_turbine.outputs.stagnation_temperature
    low_pressure_turbine.inputs.stagnation_pressure        = high_pressure_turbine.outputs.stagnation_pressure
    
    #link the low pressure turbine to the low_pressure_compresor
    low_pressure_turbine.inputs.compressor                 = low_pressure_compressor.outputs
    
    #link the low pressure turbine to the combustor
    low_pressure_turbine.inputs.fuel_to_air_ratio          = combustor.outputs.fuel_to_air_ratio
    
    #flow through the low pressure turbine
    low_pressure_turbine.inputs.bypass_ratio = 0.0
    low_pressure_turbine.inputs.fan = Data()
    low_pressure_turbine.inputs.fan.work_done = 0.0    
    low_pressure_turbine(conditions)
    
    #link the core nozzle to the low pressure turbine
    core_nozzle.inputs.stagnation_temperature              = low_pressure_turbine.outputs.stagnation_temperature
    core_nozzle.inputs.stagnation_pressure                 = low_pressure_turbine.outputs.stagnation_pressure
    
    #flow through the core nozzle
    core_nozzle(conditions)

    # compute the thrust using the thrust component
    #link the thrust component to the core nozzle
    thrust.inputs.core_exit_velocity                       = core_nozzle.outputs.velocity
    thrust.inputs.core_area_ratio                          = core_nozzle.outputs.area_ratio
    thrust.inputs.core_nozzle                              = core_nozzle.outputs
    
    #link the thrust component to the combustor
    thrust.inputs.fuel_to_air_ratio                        = combustor.outputs.fuel_to_air_ratio
    
    #link the thrust component to the low pressure compressor 
    thrust.inputs.stag_temp_lpt_exit                       = low_pressure_compressor.outputs.stagnation_temperature
    thrust.inputs.stag_press_lpt_exit                      = low_pressure_compressor.outputs.stagnation_pressure
    thrust.inputs.number_of_engines                        = number_of_engines
    thrust.inputs.total_temperature_reference              = low_pressure_compressor.outputs.stagnation_temperature
    thrust.inputs.total_pressure_reference                 = low_pressure_compressor.outputs.stagnation_pressure

    #compute the thrust
    thrust.inputs.fan_nozzle = Data()
    thrust.inputs.fan_nozzle.velocity = 0.0
    thrust.inputs.fan_nozzle.area_ratio = 0.0
    thrust.inputs.fan_nozzle.static_pressure = 0.0
    thrust.inputs.bypass_ratio = 0.0
    thrust.inputs.flow_through_core                        =  1.0 #scaled constant to turn on core thrust computation
    thrust.inputs.flow_through_fan                         =  0.0 #scaled constant to turn on fan thrust computation     
    thrust.size(conditions)
    
    #update the design thrust value
    turbojet.design_thrust = thrust.total_design

    #compute the sls_thrust
    #call the atmospheric model to get the conditions at the specified altitude
    atmosphere_sls = SUAVE.Analyses.Atmospheric.US_Standard_1976()
    atmo_data = atmosphere_sls.compute_values(0.0,0.0)
    
    p   = atmo_data.pressure          
    T   = atmo_data.temperature       
    rho = atmo_data.density          
    a   = atmo_data.speed_of_sound    
    mu  = atmo_data.dynamic_viscosity   

    # setup conditions
    conditions_sls = SUAVE.Analyses.Mission.Segments.Conditions.Aerodynamics()            

    # freestream conditions    
    conditions_sls.freestream.altitude                    = np.atleast_1d(0.)
    conditions_sls.freestream.mach_number                 = np.atleast_1d(0.01)
    conditions_sls.freestream.pressure                    = np.atleast_1d(p)
    conditions_sls.freestream.temperature                 = np.atleast_1d(T)
    conditions_sls.freestream.density                     = np.atleast_1d(rho)
    conditions_sls.freestream.dynamic_viscosity           = np.atleast_1d(mu)
    conditions_sls.freestream.gravity                     = np.atleast_1d(9.81)
    conditions_sls.freestream.isentropic_expansion_factor = np.atleast_1d(1.4)
    conditions_sls.freestream.Cp                          = 1.4*(p/(rho*T))/(1.4-1)
    conditions_sls.freestream.R                           = p/(rho*T)
    conditions_sls.freestream.speed_of_sound              = np.atleast_1d(a)
    conditions_sls.freestream.velocity                    = np.atleast_1d(a*0.01)
    
    # propulsion conditions
    conditions_sls.propulsion.throttle           =  np.atleast_1d(1.0)    
    
    state_sls                       = Data()
    state_sls.numerics              = Data()
    state_sls.conditions            = conditions_sls   
    results_sls                     = turbojet.evaluate_thrust(state_sls)
    turbojet.sealevel_static_thrust = results_sls.thrust_force_vector[0,0] / number_of_engines

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

#.........这里部分代码省略.........
    inlet_nozzle              = ducted_fan.inlet_nozzle
    fan                       = ducted_fan.fan
    fan_nozzle                = ducted_fan.fan_nozzle
    thrust                    = ducted_fan.thrust
    
    bypass_ratio              = ducted_fan.bypass_ratio #0
    number_of_engines         = ducted_fan.number_of_engines
    
    #Creating the network by manually linking the different components
    
    #set the working fluid to determine the fluid properties
    ram.inputs.working_fluid = ducted_fan.working_fluid
    
    #Flow through the ram , this computes the necessary flow quantities and stores it into conditions
    ram(conditions)

    #link inlet nozzle to ram 
    inlet_nozzle.inputs = ram.outputs
    
    #Flow through the inlet nozzle
    inlet_nozzle(conditions)
        
    #Link the fan to the inlet nozzle
    fan.inputs = inlet_nozzle.outputs
    
    #flow through the fan
    fan(conditions)        
    
    #link the dan nozzle to the fan
    fan_nozzle.inputs =  fan.outputs
    
    # flow through the fan nozzle
    fan_nozzle(conditions)
    
    # compute the thrust using the thrust component
    
    #link the thrust component to the fan nozzle
    thrust.inputs.fan_exit_velocity                        = fan_nozzle.outputs.velocity
    thrust.inputs.fan_area_ratio                           = fan_nozzle.outputs.area_ratio
    thrust.inputs.fan_nozzle                               = fan_nozzle.outputs
    thrust.inputs.number_of_engines                        = number_of_engines
    thrust.inputs.bypass_ratio                             = bypass_ratio
    thrust.inputs.total_temperature_reference              = fan_nozzle.outputs.stagnation_temperature
    thrust.inputs.total_pressure_reference                 = fan_nozzle.outputs.stagnation_pressure
    thrust.inputs.flow_through_core                        = 0.
    thrust.inputs.flow_through_fan                         = 1.
    
    #nonexistant components used to run thrust
    thrust.inputs.core_exit_velocity                       = 0.
    thrust.inputs.core_area_ratio                          = 0.
    thrust.inputs.core_nozzle                              = Data()
    thrust.inputs.core_nozzle.velocity                     = 0.
    thrust.inputs.core_nozzle.area_ratio                   = 0.
    thrust.inputs.core_nozzle.static_pressure              = 0.                                                                                                                
    
    #compute the trust
    thrust.size(conditions)
    mass_flow  = thrust.mass_flow_rate_design
    
    #update the design thrust value
    ducted_fan.design_thrust = thrust.total_design
      
    #compute the sls_thrust
    
    #call the atmospheric model to get the conditions at the specified altitude
    atmosphere_sls = SUAVE.Analyses.Atmospheric.US_Standard_1976()
    atmo_data = atmosphere_sls.compute_values(0.0,0.0)
    
    p   = atmo_data.pressure          
    T   = atmo_data.temperature       
    rho = atmo_data.density          
    a   = atmo_data.speed_of_sound    
    mu  = atmo_data.dynamic_viscosity   

    # setup conditions
    conditions_sls = SUAVE.Analyses.Mission.Segments.Conditions.Aerodynamics()            

    # freestream conditions
    conditions_sls.freestream.altitude                    = np.atleast_1d(0.)
    conditions_sls.freestream.mach_number                 = np.atleast_1d(0.01)
    conditions_sls.freestream.pressure                    = np.atleast_1d(p)
    conditions_sls.freestream.temperature                 = np.atleast_1d(T)
    conditions_sls.freestream.density                     = np.atleast_1d(rho)
    conditions_sls.freestream.dynamic_viscosity           = np.atleast_1d(mu)
    conditions_sls.freestream.gravity                     = np.atleast_1d(planet.sea_level_gravity)
    conditions_sls.freestream.isentropic_expansion_factor = np.atleast_1d(ducted_fan.working_fluid.compute_gamma(T,p))
    conditions_sls.freestream.Cp                          = np.atleast_1d(ducted_fan.working_fluid.compute_cp(T,p))
    conditions_sls.freestream.R                           = np.atleast_1d(ducted_fan.working_fluid.gas_specific_constant)
    conditions_sls.freestream.speed_of_sound              = np.atleast_1d(a)
    conditions_sls.freestream.velocity                    = conditions_sls.freestream.mach_number * conditions_sls.freestream.speed_of_sound
    
    # propulsion conditions
    conditions_sls.propulsion.throttle           =  np.atleast_1d(1.0)    
    
    state_sls = Data()
    state_sls.numerics = Data()
    state_sls.conditions = conditions_sls   
    results_sls = ducted_fan.evaluate_thrust(state_sls)
    
    ducted_fan.sealevel_static_thrust = results_sls.thrust_force_vector[0,0] / number_of_engines 

# 需要导入模块: from SUAVE.Core import Data [as 别名]
# 或者: from SUAVE.Core.Data import numerics [as 别名]
def energy_network():
    
    # ------------------------------------------------------------------
    #   Evaluation Conditions
    # ------------------------------------------------------------------    
    
    # --- Conditions        
    ones_1col = np.ones([1,1])    
    
    
    
    # setup conditions
    conditions = SUAVE.Analyses.Mission.Segments.Conditions.Aerodynamics()
    '''
    conditions.frames       = Data()
    conditions.freestream   = Data()
    conditions.aerodynamics = Data()
    conditions.propulsion   = Data()
    conditions.weights      = Data()
    conditions.energies     = Data()
    '''
  #  self.conditions = conditions
    

    # freestream conditions
    conditions.freestream.mach_number        = ones_1col*0.8
    conditions.freestream.pressure           = ones_1col*20000.
    conditions.freestream.temperature        = ones_1col*215.
    conditions.freestream.density            = ones_1col*0.8

    conditions.freestream.dynamic_viscosity          = ones_1col* 0.000001475
    conditions.freestream.altitude           = ones_1col* 10.
    conditions.freestream.gravity            = ones_1col*9.81
    conditions.freestream.gamma              = ones_1col*1.4
    conditions.freestream.Cp                 = 1.4*287.87/(1.4-1)
    conditions.freestream.R                  = 287.87
    conditions.M = conditions.freestream.mach_number 
    conditions.T = conditions.freestream.temperature
    conditions.p = conditions.freestream.pressure
    conditions.freestream.speed_of_sound     = ones_1col* np.sqrt(conditions.freestream.Cp/(conditions.freestream.Cp-conditions.freestream.R)*conditions.freestream.R*conditions.freestream.temperature) #300.
    conditions.freestream.velocity           = conditions.M * conditions.freestream.speed_of_sound
    conditions.velocity = conditions.M * conditions.freestream.speed_of_sound
    conditions.q = 0.5*conditions.freestream.density*conditions.velocity**2
    conditions.g0 = conditions.freestream.gravity
    
    # propulsion conditions
    conditions.propulsion.throttle           =  ones_1col*1.0

    
    
    # ------------------------------------------------------------------
    #   Design/sizing conditions
    # ------------------------------------------------------------------    
    
    
    # --- Conditions        
    ones_1col = np.ones([1,1])    
    
    
    
    # setup conditions
    conditions_sizing = SUAVE.Analyses.Mission.Segments.Conditions.Aerodynamics()
    '''
    conditions_sizing.frames       = Data()
    conditions_sizing.freestream   = Data()
    conditions_sizing.aerodynamics = Data()
    conditions_sizing.propulsion   = Data()
    conditions_sizing.weights      = Data()
    conditions_sizing.energies     = Data()
    '''
  #  self.conditions = conditions
    

    # freestream conditions
    conditions_sizing.freestream.mach_number        = ones_1col*0.8
    conditions_sizing.freestream.pressure           = ones_1col*26499.73156529
    conditions_sizing.freestream.temperature        = ones_1col*223.25186491
    conditions_sizing.freestream.density            = ones_1col*0.41350854

    conditions_sizing.freestream.dynamic_viscosity  = ones_1col* 1.45766126e-05 #*1.789*10**(-5)
    conditions_sizing.freestream.altitude           = ones_1col* 10000. #* 0.5

    conditions_sizing.freestream.gravity            = ones_1col*9.81
    conditions_sizing.freestream.gamma              = ones_1col*1.4
    conditions_sizing.freestream.Cp                 = 1.4*287.87/(1.4-1)
    conditions_sizing.freestream.R                  = 287.87
    conditions_sizing.freestream.speed_of_sound     = 299.53150968
    conditions_sizing.freestream.velocity           = conditions_sizing.freestream.mach_number * conditions_sizing.freestream.speed_of_sound
    
    # propulsion conditions
    conditions_sizing.propulsion.throttle           =  ones_1col*1.0

    state_sizing = Data()
    state_sizing.numerics = Data()
    state_sizing.conditions = conditions_sizing
    state_off_design=Data()
    state_off_design.numerics=Data()
    state_off_design.conditions=conditions


#.........这里部分代码省略.........

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

#.........这里部分代码省略.........
    
    #link inlet nozzle to ram 
    inlet_nozzle.inputs.stagnation_temperature             = ram.outputs.stagnation_temperature #conditions.freestream.stagnation_temperature
    inlet_nozzle.inputs.stagnation_pressure                = ram.outputs.stagnation_pressure #conditions.freestream.stagnation_pressure
    
    #Flow through the inlet nozzle
    inlet_nozzle(conditions)
        
    #Link the fan to the inlet nozzle
    fan.inputs.stagnation_temperature                      = inlet_nozzle.outputs.stagnation_temperature
    fan.inputs.stagnation_pressure                         = inlet_nozzle.outputs.stagnation_pressure
    
    #flow through the fan
    fan(conditions)
    
    
           
    

    #link the dan nozzle to the fan
    fan_nozzle.inputs.stagnation_temperature               = fan.outputs.stagnation_temperature
    fan_nozzle.inputs.stagnation_pressure                  = fan.outputs.stagnation_pressure
    
     # flow through the fan nozzle
    fan_nozzle(conditions)
    
    # compute the thrust using the thrust component
    
    
    
    #link the thrust component to the fan nozzle
    thrust.inputs.fan_exit_velocity                        = fan_nozzle.outputs.velocity
    thrust.inputs.fan_area_ratio                           = fan_nozzle.outputs.area_ratio
    thrust.inputs.fan_nozzle                               = fan_nozzle.outputs
    thrust.inputs.number_of_engines                        = number_of_engines
    thrust.inputs.bypass_ratio                             = bypass_ratio
    thrust.inputs.total_temperature_reference              = fan_nozzle.outputs.stagnation_temperature
    thrust.inputs.total_pressure_reference                 = fan_nozzle.outputs.stagnation_pressure
    thrust.inputs.flow_through_core                        = 0.
    thrust.inputs.flow_through_fan                         = 1.
    
    #nonexistant components used to run thrust
    thrust.inputs.core_exit_velocity                       = 0.
    thrust.inputs.core_area_ratio                          = 0.
    thrust.inputs.core_nozzle                              = Data()
    thrust.inputs.core_nozzle.velocity                     = 0.
    thrust.inputs.core_nozzle.area_ratio                   = 0.
    thrust.inputs.core_nozzle.static_pressure              = 0.                                                                                                                
    
    
    #compute the trust
    thrust.size(conditions)
    mass_flow  = thrust.mass_flow_rate_design
    

    #update the design thrust value
    ducted_fan.design_thrust = thrust.total_design
    
    
    #compute the sls_thrust
    
    #call the atmospheric model to get the conditions at the specified altitude
    atmosphere_sls = SUAVE.Analyses.Atmospheric.US_Standard_1976()
    p,T,rho,a,mu = atmosphere_sls.compute_values(0.0,0.0)

    # setup conditions
    conditions_sls = SUAVE.Analyses.Mission.Segments.Conditions.Aerodynamics()            



    # freestream conditions
    
    conditions_sls.freestream.altitude           = np.atleast_1d(0.)
    conditions_sls.freestream.mach_number        = np.atleast_1d(0.01)
    
    conditions_sls.freestream.pressure           = np.atleast_1d(p)
    conditions_sls.freestream.temperature        = np.atleast_1d(T)
    conditions_sls.freestream.density            = np.atleast_1d(rho)
    conditions_sls.freestream.dynamic_viscosity  = np.atleast_1d(mu)
    conditions_sls.freestream.gravity            = np.atleast_1d(9.81)
    conditions_sls.freestream.gamma              = np.atleast_1d(1.4)
    conditions_sls.freestream.Cp                 = 1.4*287.87/(1.4-1)
    conditions_sls.freestream.R                  = 287.87
    conditions_sls.freestream.speed_of_sound     = np.atleast_1d(a)
    conditions_sls.freestream.velocity           = conditions_sls.freestream.mach_number * conditions_sls.freestream.speed_of_sound
    
    # propulsion conditions
    conditions_sls.propulsion.throttle           =  np.atleast_1d(1.0)    
    
    state_sls = Data()
    state_sls.numerics = Data()
    state_sls.conditions = conditions_sls   
    results_sls = ducted_fan.evaluate_thrust(state_sls)
    
    ducted_fan.sealevel_static_thrust = results_sls.thrust_force_vector[0,0] / number_of_engines
   
    
    #determine geometry
    
    

# 需要导入模块: from SUAVE.Core import Data [as 别名]
# 或者: from SUAVE.Core.Data import numerics [as 别名]
def main():
   
    #------------------------------------------------------------------
    # Propulsor
    #------------------------------------------------------------------

    # build network
    net = Solar_Low_Fidelity()
    net.number_of_engines = 1.
    net.nacelle_diameter  = 0.05
    net.areas             = Data()
    net.areas.wetted      = 0.01*(2*np.pi*0.01/2)
    net.engine_length     = 0.01

    # 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.ratio                = 0.9
    panel.area                 = 1.0 * panel.ratio 
    panel.efficiency           = 0.25
    panel.mass_properties.mass = panel.area*(0.60 * Units.kg)
    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
    prop = SUAVE.Components.Energy.Converters.Propeller_Lo_Fid()
    prop.propulsive_efficiency = 0.825
    net.propeller        = prop
    
    # Component 4 the Motor
    motor = SUAVE.Components.Energy.Converters.Motor_Lo_Fid()
    kv                         = 800. * Units['rpm/volt'] # RPM/volt is standard
    motor                      = size_from_kv(motor, kv)    
    motor.gear_ratio           = 1. # Gear ratio, no gearbox
    motor.gearbox_efficiency   = 1. # Gear box efficiency, no gearbox
    motor.motor_efficiency     = 0.825;
    net.motor                  = motor    

    # Component 6 the Payload
    payload = SUAVE.Components.Energy.Peripherals.Payload()
    payload.power_draw           = 0. #Watts 
    payload.mass_properties.mass = 0.0 * Units.kg
    net.payload                  = payload

    # Component 7 the Avionics
    avionics = SUAVE.Components.Energy.Peripherals.Avionics()
    avionics.power_draw = 10. #Watts  
    net.avionics        = avionics      

    # Component 8 the Battery
    bat = SUAVE.Components.Energy.Storages.Batteries.Constant_Mass.Lithium_Ion()
    bat.mass_properties.mass = 5.0  * Units.kg
    bat.specific_energy      = 250. *Units.Wh/Units.kg
    bat.resistance           = 0.003
    bat.iters                = 0
    initialize_from_mass(bat,bat.mass_properties.mass)
    net.battery              = bat

    #Component 9 the system logic controller and MPPT
    logic = SUAVE.Components.Energy.Distributors.Solar_Logic()
    logic.system_voltage  = 18.5
    logic.MPPT_efficiency = 0.95
    net.solar_logic       = logic
    
    # Setup the conditions to run the network
    state            = Data()
    state.conditions = SUAVE.Analyses.Mission.Segments.Conditions.Aerodynamics()
    state.numerics   = SUAVE.Analyses.Mission.Segments.Conditions.Numerics()
    
    conditions = state.conditions
    numerics   = state.numerics
    
    # Calculate atmospheric properties
    atmosphere = SUAVE.Analyses.Atmospheric.US_Standard_1976()
    atmosphere_conditions =  atmosphere.compute_values(1000.*Units.ft)
    
    rho = atmosphere_conditions.density[0,:]
    a   = atmosphere_conditions.speed_of_sound[0,:]
    mu  = atmosphere_conditions.dynamic_viscosity[0,:]
    T   = atmosphere_conditions.temperature[0,:]

    conditions.propulsion.throttle            = np.array([[1.0],[1.0]])
    conditions.freestream.velocity            = np.array([[1.0],[1.0]])
    conditions.freestream.density             = np.array([rho,rho])
    conditions.freestream.dynamic_viscosity   = np.array([mu, mu])
    conditions.freestream.speed_of_sound      = np.array([a, a])
    conditions.freestream.altitude            = np.array([[1000.0],[1000.0]])
    conditions.propulsion.battery_energy      = bat.max_energy*np.ones_like(conditions.freestream.altitude)
    conditions.frames.body.inertial_rotations = np.zeros([2,3])
    conditions.frames.inertial.time           = np.array([[0.0],[1.0]])
    numerics.time.integrate                   = np.array([[0, 0],[0, 1]])
    numerics.time.differentiate               = np.array([[0, 0],[0, 1]])
    conditions.frames.planet.start_time       = time.strptime("Sat, Jun 21 06:00:00  2014", "%a, %b %d %H:%M:%S %Y",) 
#.........这里部分代码省略.........

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

#.........这里部分代码省略.........
            conditions.freestream.Cp                          = np.atleast_1d(scramjet.working_fluid.compute_cp(T,p))
            conditions.freestream.R                           = np.atleast_1d(scramjet.working_fluid.gas_specific_constant)
            conditions.freestream.speed_of_sound              = np.atleast_1d(a)
            conditions.freestream.velocity                    = np.atleast_1d(a*mach_number)
            
            # propulsion conditions
            conditions.propulsion.throttle           =  np.atleast_1d(1.0)
    
    ram                       = scramjet.ram
    inlet_nozzle              = scramjet.inlet_nozzle
    combustor                 = scramjet.combustor
    core_nozzle               = scramjet.core_nozzle
    thrust                    = scramjet.thrust
    number_of_engines         = scramjet.number_of_engines
    
    #Creating the network by manually linking the different components
    #set the working fluid to determine the fluid properties
    ram.inputs.working_fluid       = scramjet.working_fluid
    
    #Flow through the ram
    ram(conditions)
    
    #link inlet nozzle to ram 
    inlet_nozzle.inputs             = ram.outputs
    
    #Flow through the inlet nozzle
    inlet_nozzle.compute_scramjet(conditions)

    #link the combustor to the inlet nozzle
    combustor.inputs.stagnation_temperature                = inlet_nozzle.outputs.stagnation_temperature 
    combustor.inputs.stagnation_pressure                   = inlet_nozzle.outputs.stagnation_pressure 
    combustor.inputs.inlet_nozzle                          = inlet_nozzle.outputs
        
    #flow through the high pressor comprresor
    combustor.compute_supersonic_combustion(conditions)
    
    #link the core nozzle to the combustor
    core_nozzle.inputs              = combustor.outputs
    
    #flow through the core nozzle
    core_nozzle.compute_scramjet(conditions)

    #link the thrust component to the core nozzle
    thrust.inputs.core_nozzle                              = core_nozzle.outputs
    thrust.inputs.number_of_engines                        = number_of_engines
    thrust.inputs.total_temperature_reference              = core_nozzle.outputs.stagnation_temperature
    thrust.inputs.total_pressure_reference                 = core_nozzle.outputs.stagnation_pressure
    
    #link the thrust component to the combustor
    thrust.inputs.fuel_to_air_ratio = combustor.outputs.fuel_to_air_ratio
    
    #compute the thrust
    thrust.inputs.fan_nozzle                               = Data()
    thrust.inputs.fan_nozzle.velocity                      = 0.0
    thrust.inputs.fan_nozzle.area_ratio                    = 0.0
    thrust.inputs.fan_nozzle.static_pressure               = 0.0
    thrust.inputs.bypass_ratio                             = 0.0
    thrust.inputs.flow_through_core                        = 1.0 #scaled constant to turn on core thrust computation
    thrust.inputs.flow_through_fan                         = 0.0 #scaled constant to turn on fan thrust computation     
    thrust.size_stream_thrust(conditions)
    
    #update the design thrust value
    scramjet.design_thrust = thrust.total_design

    #compute the sls_thrust
    #call the atmospheric model to get the conditions at the specified altitude
    atmosphere_sls  = SUAVE.Analyses.Atmospheric.US_Standard_1976()
    atmo_data       = atmosphere_sls.compute_values(0.0,0.0)
    
    p   = atmo_data.pressure          
    T   = atmo_data.temperature       
    rho = atmo_data.density          
    a   = atmo_data.speed_of_sound    
    mu  = atmo_data.dynamic_viscosity   

    # setup conditions
    conditions_sls = SUAVE.Analyses.Mission.Segments.Conditions.Aerodynamics()            

    # freestream conditions    
    conditions_sls.freestream.altitude                    = np.atleast_1d(0.)
    conditions_sls.freestream.mach_number                 = np.atleast_1d(0.01)
    conditions_sls.freestream.pressure                    = np.atleast_1d(p)
    conditions_sls.freestream.temperature                 = np.atleast_1d(T)
    conditions_sls.freestream.density                     = np.atleast_1d(rho)
    conditions_sls.freestream.dynamic_viscosity           = np.atleast_1d(mu)
    conditions_sls.freestream.gravity                     = np.atleast_1d(planet.sea_level_gravity)
    conditions_sls.freestream.isentropic_expansion_factor = np.atleast_1d(scramjet.working_fluid.compute_gamma(T,p))
    conditions_sls.freestream.Cp                          = np.atleast_1d(scramjet.working_fluid.compute_cp(T,p))
    conditions_sls.freestream.R                           = np.atleast_1d(scramjet.working_fluid.gas_specific_constant)
    conditions_sls.freestream.speed_of_sound              = np.atleast_1d(a)
    conditions_sls.freestream.velocity                    = np.atleast_1d(a*0.01)
    
    # propulsion conditions
    conditions_sls.propulsion.throttle           =  np.atleast_1d(1.0)    
    
    state_sls            = Data()
    state_sls.numerics   = Data()
    state_sls.conditions = conditions_sls   
    results_sls          = scramjet.evaluate_thrust(state_sls)
    scramjet.sealevel_static_thrust = results_sls.thrust_force_vector[0,0] / number_of_engines

# 需要导入模块: from SUAVE.Core import Data [as 别名]
# 或者: from SUAVE.Core.Data import numerics [as 别名]
def energy_network():
    
    # ------------------------------------------------------------------
    #   Evaluation Conditions
    # ------------------------------------------------------------------    
    
    # --- Conditions        
    ones_1col = np.ones([1,1])    
    
    # setup conditions
    conditions = SUAVE.Analyses.Mission.Segments.Conditions.Aerodynamics()
    
    # freestream conditions
    EVAL                             = conditions.freestream
    EVAL.mach_number                 = ones_1col*1.5
    conditions.M                     = EVAL.mach_number
    EVAL.altitude                    = ones_1col*10000.
    
    atmosphere                       = SUAVE.Analyses.Atmospheric.US_Standard_1976()
    atmo_data                        = atmosphere.compute_values(EVAL.altitude,0,True) 
    working_fluid                    = SUAVE.Attributes.Gases.Air()
    
    EVAL.pressure                    = ones_1col*atmo_data.pressure
    EVAL.temperature                 = ones_1col*atmo_data.temperature
    EVAL.density                     = ones_1col*atmo_data.density
    EVAL.dynamic_viscosity           = ones_1col* atmo_data.dynamic_viscosity
    EVAL.gravity                     = ones_1col*9.81
    EVAL.isentropic_expansion_factor = working_fluid.compute_gamma(EVAL.temperature,EVAL.pressure)
    EVAL.Cp                          = working_fluid.compute_cp(EVAL.temperature,EVAL.pressure)                                                                               
    EVAL.R                           = working_fluid.gas_specific_constant
    EVAL.speed_of_sound              = ones_1col* atmo_data.speed_of_sound
    EVAL.velocity                    = conditions.M * EVAL.speed_of_sound
    conditions.velocity              = conditions.M * EVAL.speed_of_sound
    conditions.q                     = 0.5*EVAL.density*conditions.velocity**2
    conditions.g0                    = EVAL.gravity
    
    # propulsion conditions
    conditions.propulsion.throttle   =  ones_1col*1.0

    # ------------------------------------------------------------------
    #   Design/sizing conditions
    # ------------------------------------------------------------------    
    
    # --- Conditions        
    ones_1col = np.ones([1,1])    
    
    # setup conditions
    conditions_sizing = SUAVE.Analyses.Mission.Segments.Conditions.Aerodynamics()

    # freestream conditions
    SIZE                             = conditions_sizing.freestream
    SIZE.mach_number                 = ones_1col*2.5
    conditions_sizing.M              = SIZE.mach_number
    SIZE.altitude                    = ones_1col*10000.  
    
    atmosphere                       = SUAVE.Analyses.Atmospheric.US_Standard_1976()
    atmo_data                        = atmosphere.compute_values(SIZE.altitude,0,True) 
    working_fluid                    = SUAVE.Attributes.Gases.Air()    

    SIZE.pressure                    = ones_1col*atmo_data.pressure
    SIZE.temperature                 = ones_1col*atmo_data.temperature
    SIZE.density                     = ones_1col*atmo_data.density
    SIZE.dynamic_viscosity           = ones_1col*atmo_data.dynamic_viscosity
    SIZE.gravity                     = ones_1col*9.81
    SIZE.isentropic_expansion_factor = working_fluid.compute_gamma(SIZE.temperature,SIZE.pressure)
    SIZE.Cp                          = working_fluid.compute_cp(SIZE.temperature,SIZE.pressure)                                                                               
    SIZE.R                           = working_fluid.gas_specific_constant
    SIZE.speed_of_sound              = ones_1col * atmo_data.speed_of_sound
    SIZE.velocity                    = conditions_sizing.M * SIZE.speed_of_sound
    conditions_sizing.velocity       = conditions_sizing.M * SIZE.speed_of_sound
    conditions_sizing.q              = 0.5*SIZE.density*conditions_sizing.velocity**2
    conditions_sizing.g0             = SIZE.gravity
    
    # propulsion conditions
    conditions_sizing.propulsion.throttle = ones_1col*1.0

    state_sizing = Data()
    state_sizing.numerics = Data()
    state_sizing.conditions = conditions_sizing
    state_off_design=Data()
    state_off_design.numerics=Data()
    state_off_design.conditions=conditions

    # ------------------------------------------------------------------
    #  Ramjet Network
    # ------------------------------------------------------------------    
    
    # instantiate the ramjet network
    ramjet = SUAVE.Components.Energy.Networks.Ramjet()
    ramjet.tag = 'ramjet'
    
    # setup
    ramjet.number_of_engines = 2.0
    ramjet.engine_length     = 6.0
    ramjet.nacelle_diameter  = 1.3 * Units.meter
    ramjet.inlet_diameter    = 1.1 * Units.meter
    
    # working fluid
    ramjet.working_fluid = SUAVE.Attributes.Gases.Air()

#.........这里部分代码省略.........