How to use the gpkit.Vectorize function in gpkit

setup method for the engine model
        """
        self.setvals(eng)
        self.compressor = Compressor()
        self.combustor = Combustor()
        self.turbine = Turbine()
        self.fanmap = FanMap()
        self.lpcmap = LPCMap()
        self.hpcmap = HPCMap()
        self.thrust = Thrust()
        self.sizing = Sizing()
        self.constants = EngineConstants()
        self.state = state

        if Nfleet != 0:
            with Vectorize(Nfleet):
                with Vectorize(N):
                    #-------------------Specified Thrust or Tt4-----------------------
                    self.engineP = self.dynamic(self.state, res7)
                    if res7 == 0:
                        #variables for the thrust constraint
                        Fspec = Variable('F_{spec}', 'N', 'Specified Total Thrust')
                    if res7 == 1:
                        Tt4spec = Variable('T_{t_{4spec}}', 'K', 'Specified Combustor Exit (Station 4) Stagnation Temperature')
                    

        else:
            with Vectorize(N):
                #-------------------Specified Thrust or Tt4-----------------------
                self.engineP = self.dynamic(self.state, res7)
                if res7 == 0:
                    #variables for the thrust constraint

Ncruise = 2

        eng = 0
        
        # vectorize
        with Vectorize(Nclimb + Ncruise):
            enginestate = FlightState()

        # build required submodels
        aircraft = Aircraft(Nclimb, Ncruise, enginestate, eng)

        # vectorize
        with Vectorize(Nclimb):
            climb = ClimbSegment(aircraft)

        with Vectorize(Ncruise):
            cruise = CruiseSegment(aircraft)

        statelinking = StateLinking(climb.state, cruise.state, enginestate, Nclimb, Ncruise)

        # declare new variables
        W_ftotal = Variable('W_{f_{total}}', 'lbf', 'Total Fuel Weight')
        W_fclimb = Variable('W_{f_{climb}}', 'lbf',
                            'Fuel Weight Burned in Climb')
        W_fcruise = Variable('W_{f_{cruise}}', 'lbf',
                             'Fuel Weight Burned in Cruise')
        W_total = Variable('W_{total}', 'lbf', 'Total Aircraft Weight')
        W_dry = Variable('W_{dry}', 'lbf', 'Dry Aircraft Weight')
        CruiseAlt = Variable('CruiseAlt', 'ft', 'Cruise Altitude [feet]')
        ReqRng = Variable('ReqRng', 'nautical_miles', 'Required Cruise Range')

        h = climb.state['h']

def setup(self,static,  state, N=5):
        with Vectorize(N):
            blade = BladeElementPerf(static, state)


        constraints = [blade.dr == static.R/(N),
                        blade.omega == omega,
                        blade.r[0] == static.R/(2.*N)]

        for n in range(1,N):
            constraints += [TCS([blade.r[n] >= blade.r[n-1] + static.R/N]),
                            blade.eta_i[n] == blade.eta_i[n-1],
                            ]

        constraints += [TCS([Q >= sum(blade.dQ)]),
                        eta == state.V*T/(omega*Q),
                        blade.M[-1] <= Mtip,
                        static.T_m >= T,

if conventional:
            wingengine = True; tube = True;

        # Multimission?
        if Nmission == 1:
            multimission = False
        else:
            multimission = True

        # Defining fitDrag, boolean describing whether or not to use XFOIL tail drag fits
        # False uses TASOPT tail drag model. Currently on.
        fitDrag = True

        # vectorize
        with Vectorize(Nmission):
             with Vectorize(Nclimb + Ncruise):
                 self.flightstate = flightstate = FlightState()

        # Build required submodels
        self.aircraft = aircraft = Aircraft(Nclimb, Ncruise, flightstate, eng, fitDrag, BLI, Nmission)
        self.aircraft.config = config

        # Vectorize dynamic variables
        with Vectorize(Nmission):
             with Vectorize(Nclimb+Ncruise):
                 self.flight = flight = FlightSegment(aircraft, flightstate, Nclimb, Ncruise)

        # Declare Mission variables
        if multimission:
             with Vectorize(Nmission):
                  CruiseAlt = Variable('CruiseAlt', 'ft', 'Cruise Altitude [feet]')
                  ReqRng = Variable('R_{req}', 'nautical_miles', 'Required Cruise Range')

def __init__(self, ac, substitutions = None, **kwargs):
        #define the number of each flight segment
        Nclimb = 2
        Ncruise = 2

        #Vectorize
        with Vectorize(Nclimb):
            climb = ClimbSegment(ac)

        with Vectorize(Ncruise):
            cruise = CruiseSegment(ac)

        #declare new variables
        W_ftotal = Variable('W_{f_{total}}', 'N', 'Total Fuel Weight')
        W_fclimb = Variable('W_{f_{climb}}', 'N', 'Fuel Weight Burned in Climb')
        W_fcruise = Variable('W_{f_{cruise}}', 'N', 'Fuel Weight Burned in Cruise')
        W_total = Variable('W_{total}', 'N', 'Total Aircraft Weight')
        CruiseAlt = Variable('CruiseAlt', 'ft', 'Cruise Altitude [feet]')
        ReqRng = Variable('ReqRng', 'nautical_miles', 'Required Cruise Range')

        xcg     = Variable('x_{CG}', 'm', 'CG location')

        h = climb['h']
        hftClimb = climb['hft']
        dhft = climb['dhft']
        hftCruise = cruise['hft']

"""
        self.setvals(eng)
        self.compressor = Compressor()
        self.combustor = Combustor()
        self.turbine = Turbine()
        self.fanmap = FanMap()
        self.lpcmap = LPCMap()
        self.hpcmap = HPCMap()
        self.thrust = Thrust()
        self.sizing = Sizing()
        self.constants = EngineConstants()
        self.state = state

        if Nfleet != 0:
            with Vectorize(Nfleet):
                with Vectorize(N):
                    #-------------------Specified Thrust or Tt4-----------------------
                    self.engineP = self.dynamic(self.state, res7)
                    if res7 == 0:
                        #variables for the thrust constraint
                        Fspec = Variable('F_{spec}', 'N', 'Specified Total Thrust')
                    if res7 == 1:
                        Tt4spec = Variable('T_{t_{4spec}}', 'K', 'Specified Combustor Exit (Station 4) Stagnation Temperature')
                    

        else:
            with Vectorize(N):
                #-------------------Specified Thrust or Tt4-----------------------
                self.engineP = self.dynamic(self.state, res7)
                if res7 == 0:
                    #variables for the thrust constraint
                    Fspec = Variable('F_{spec}', 'N', 'Specified Total Thrust')

def setup(self, **kwargs):
        # define the number of each flight segment
        Nclimb = 2
        Ncruise = 2

        eng = 0
        
        # vectorize
        with Vectorize(Nclimb + Ncruise):
            enginestate = FlightState()

        # build required submodels
        aircraft = Aircraft(Nclimb, Ncruise, enginestate, eng)

        # vectorize
        with Vectorize(Nclimb):
            climb = ClimbSegment(aircraft)

        with Vectorize(Ncruise):
            cruise = CruiseSegment(aircraft)

        statelinking = StateLinking(climb.state, cruise.state, enginestate, Nclimb, Ncruise)

        # declare new variables
        W_ftotal = Variable('W_{f_{total}}', 'lbf', 'Total Fuel Weight')

def setup(self):
        #define the number of each flight segment
        Nclimb = 2
        Ncruise = 2

        #build required submodels
        ac = Aircraft()

        #vectorize
        with Vectorize(Nclimb):
            climb = ClimbSegment(ac)

        with Vectorize(Ncruise):
            cruise = CruiseSegment(ac)

        #declare new variables
        W_ftotal = Variable('W_{f_{total}}', 'N', 'Total Fuel Weight')
        W_fclimb = Variable('W_{f_{climb}}', 'N', 'Fuel Weight Burned in Climb')
        W_fcruise = Variable('W_{f_{cruise}}', 'N', 'Fuel Weight Burned in Cruise')
        W_total = Variable('W_{total}', 'N', 'Total Aircraft Weight')
        CruiseAlt = Variable('CruiseAlt', 'ft', 'Cruise Altitude [feet]')
        ReqRng = Variable('ReqRng', 'nautical_miles', 'Required Cruise Range')

        h = climb.state['h']
        hftClimb = climb.state['hft']
        dhft = climb.climbP['dhft']
        hftCruise = cruise.state['hft']

        #make overall constraints

M4a = .1025
        Mexit = 1
        M0 = .4596

##        climb = [
##            engine.state['P_{atm}'][1] == 72.4*units('kPa'),    #36K feet
##            engine.state["T_{atm}"][1] == 270*units('K'),
##            engine.state['M'][1] == M0,
##            engine.engineP['M_2'][1] == M2,
##            engine.engineP['M_{2.5}'][1] == M25,
##            ]

        return rotation, toclimb, cruise

if __name__ == "__main__":
    with Vectorize(3):
        state = TestState()

    engine = Engine(0, True, 3, state)
    mission = TestMission(engine)

    M4a = .1025
    fan = 1.685
    lpc  = 4.744
    hpc = 3.75
 
    substitutions = {
            '\\pi_{tn}': .989,
            '\pi_{b}': .94,
            '\pi_{d}': .998,
            '\pi_{fn}': .98,
            'T_{ref}': 288.15,

def __init__(self, subs = None, **kwargs):
        #define the number of each flight segment
        Nclimb = 2
        Ncruise = 2

        #build required submodels
        ac = Aircraft()

        #Vectorize
        with Vectorize(Nclimb):
            climb = ClimbSegment(ac)

        with Vectorize(Ncruise):
            cruise = CruiseSegment(ac)

        #declare new variables
        W_ftotal = Variable('W_{f_{total}}', 'N', 'Total Fuel Weight')
        W_fclimb = Variable('W_{f_{climb}}', 'N', 'Fuel Weight Burned in Climb')
        W_fcruise = Variable('W_{f_{cruise}}', 'N', 'Fuel Weight Burned in Cruise')
        W_total = Variable('W_{total}', 'N', 'Total Aircraft Weight')
        CruiseAlt = Variable('CruiseAlt', 'ft', 'Cruise Altitude [feet]')
        ReqRng = Variable('ReqRng', 'nautical_miles', 'Required Cruise Range')

        h = climb.state.alt['h']
        hftClimb = climb.state.alt['hft']
        dhft = climb.climbP['dhft']
        hftCruise = cruise.state.alt['hft']

        #make overall constraints