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gpkit.tests.from_paths.TestFiles.test_model_print_perf_py_mosek_cli (from gpkit.tests.from_paths.TestFiles-20200714230033)

Failing for the past 1 build (Since Failed #446 )
Took 0.87 sec.

Error Message

Invalid ConstraintSet element '[<gpkit.Climb object containing 2 top-level constraint(s) and 933 variable(s)>, <gpkit.Cruise object containing 2 top-level constraint(s) and 364 variable(s)>, <gpkit.Loiter object containing 2 top-level constraint(s) and 612 variable(s)>, <gpkit.Cruise object containing 2 top-level constraint(s) and 364 variable(s)>]' <class 'list'> was between Aircraft ========  Cost ----  1  Constraints -----------   Fuselage    FuelTank     Mission1.Aircraft.Fuselage.FuelTank.W >= f*W_{fuel-tot}     \mathcal{V}/Mission1.Aircraft.Fuselage.FuelTank.m_{fac} >= W_{fuel-tot}/\rho_{fuel}     FuselageSkin     m >= S_{wet}*\rho_{kevlar}*Mission1.Aircraft.Fuselage.FuselageSkin.t     Mission1.Aircraft.Fuselage.FuselageSkin.W >= m*Mission1.Aircraft.Fuselage.FuselageSkin.g     Mission1.Aircraft.Fuselage.FuselageSkin.t >= t_{min}     Mission1.Aircraft.Fuselage.FuselageSkin.I <= PI*R^3*Mission1.Aircraft.Fuselage.FuselageSkin.t     I_G >= m*(4*R^2 + 4*R*Mission1.Aircraft.Fuselage.FuselageSkin.t + Mission1.Aircraft.Fuselage.FuselageSkin.t^2)     l_{body} = l_{body}     Mission1.Aircraft.Fuselage.FuselageSkin.E = Mission1.Aircraft.Fuselage.FuselageSkin.E    k_{body} = l_{body}/R    S_{wet} >= S_{body} + S_{nose} + S_{bulk}    S_{body} >= 2PI*R*l_{body}    S_{nose}^1.6 >= 2PI*R^2^1.6*(0.333 + 0.667*k_{nose}^1.6)    S_{bulk} >= R^2*(0.0123*k_{bulk}^2 + 1.52*k_{bulk} + 0.502)    \mathcal{V}_{body} <= PI*R^2*l_{body}    Mission1.Aircraft.Fuselage.l <= 3*R*k_{body}*k_{nose}*k_{bulk}^0.3333333333333333    Mission1.Aircraft.Fuselage.S >= PI*R^2    \mathcal{V}_{body} >= \mathcal{V}    Mission1.Aircraft.Fuselage.W/Mission1.Aircraft.Fuselage.m_{fac} >= Mission1.Aircraft.Fuselage.FuelTank.W + Mission1.Aircraft.Fuselage.FuselageSkin.W    Wing    Mission1.Aircraft.Wing.W/Mission1.Aircraft.Wing.mfac >= Mission1.Aircraft.Wing.WingSkin.W + Mission1.Aircraft.Wing.CapSpar.W + Mission1.Aircraft.Wing.WingCore.W     Planform     Mission1.Aircraft.Wing.Planform.b^2 = Mission1.Aircraft.Wing.Planform.S*Mission1.Aircraft.Wing.Planform.AR     cave[:] = cbave[:]*Mission1.Aircraft.Wing.Planform.S/Mission1.Aircraft.Wing.Planform.b     Mission1.Aircraft.Wing.Planform.croot = Mission1.Aircraft.Wing.Planform.S/Mission1.Aircraft.Wing.Planform.b*cbar[0]     Mission1.Aircraft.Wing.Planform.cmac = Mission1.Aircraft.Wing.Planform.croot*Mission1.Aircraft.Wing.Planform.cbarmac    WingSkin     Mission1.Aircraft.Wing.WingSkin.W >= CFRPFabric.rho*Mission1.Aircraft.Wing.Planform.S*2*Mission1.Aircraft.Wing.WingSkin.t*g     Mission1.Aircraft.Wing.WingSkin.t >= tmin     CFRPFabric.tau >= 1/Mission1.Aircraft.Wing.WingSkin.Jtbar/Mission1.Aircraft.Wing.Planform.croot^2/Mission1.Aircraft.Wing.WingSkin.t*Mission1.Aircraft.Wing.WingSkin.Cmw*Mission1.Aircraft.Wing.Planform.S*Mission1.Aircraft.Wing.WingSkin.rhosl*Mission1.Aircraft.Wing.WingSkin.Vne^2     CapSpar     I[:]/Mission1.Aircraft.Wing.CapSpar.mfac <= 2*w[:]*t[:]*(hin[:]/2)^2     dm[:] >= (CFRPUD.rho*2*w[:]*t[:] + 2*tshear[:]*CFRPFabric.rho*(hin[:] + 2*t[:]) + FoamHD.rho*w[:]*hin[:])*Mission1.Aircraft.Wing.Planform.b/2*deta[:]     Mission1.Aircraft.Wing.CapSpar.W >= 2*dm[:].sum()*g     w[:] <= wlim*cave[:]     cave[:]*Mission1.Aircraft.Wing.Planform.tau >= hin[:] + 2*t[:]     Sy[:]*(hin[:]/2 + t[:]) <= I[:]     tshear[:] >= tmin     WingCore     Mission1.Aircraft.Wing.WingCore.W >= 2*(g*FoamHD.rho*Mission1.Aircraft.Wing.WingCore.Abar*cave[:]^2*Mission1.Aircraft.Wing.Planform.b/2*deta[:]).sum()    mw*(1 + 2/Mission1.Aircraft.Wing.Planform.AR) >= 6.28    DF70    W_{DF70} <= Mission1.Aircraft.DF70.W/Mission1.Aircraft.DF70.m_{fac}    P_{sl-max} = P_{sl-max}    Empennage    HorizontalTail     Mission1.Aircraft.Empennage.HorizontalTail.W/Mission1.Aircraft.Empennage.HorizontalTail.mfac >= Mission1.Aircraft.Empennage.HorizontalTail.WingSkin.W + Mission1.Aircraft.Empennage.HorizontalTail.WingCore.W      Planform      Mission1.Aircraft.Empennage.HorizontalTail.Planform.b^2 = Mission1.Aircraft.Empennage.HorizontalTail.Planform.S*Mission1.Aircraft.Empennage.HorizontalTail.Planform.AR      cave[:] = cbave[:]*Mission1.Aircraft.Empennage.HorizontalTail.Planform.S/Mission1.Aircraft.Empennage.HorizontalTail.Planform.b      Mission1.Aircraft.Empennage.HorizontalTail.Planform.croot = Mission1.Aircraft.Empennage.HorizontalTail.Planform.S/Mission1.Aircraft.Empennage.HorizontalTail.Planform.b*cbar[0]      Mission1.Aircraft.Empennage.HorizontalTail.Planform.cmac = Mission1.Aircraft.Empennage.HorizontalTail.Planform.croot*Mission1.Aircraft.Empennage.HorizontalTail.Planform.cbarmac     WingSkin      Mission1.Aircraft.Empennage.HorizontalTail.WingSkin.W >= CFRPFabric.rho*Mission1.Aircraft.Empennage.HorizontalTail.Planform.S*2*Mission1.Aircraft.Empennage.HorizontalTail.WingSkin.t*g      Mission1.Aircraft.Empennage.HorizontalTail.WingSkin.t >= tmin      CFRPFabric.tau >= 1/Mission1.Aircraft.Empennage.HorizontalTail.WingSkin.Jtbar/Mission1.Aircraft.Empennage.HorizontalTail.Planform.croot^2/Mission1.Aircraft.Empennage.HorizontalTail.WingSkin.t*Mission1.Aircraft.Empennage.HorizontalTail.WingSkin.Cmw*Mission1.Aircraft.Empennage.HorizontalTail.Planform.S*Mission1.Aircraft.Empennage.HorizontalTail.WingSkin.rhosl*Mission1.Aircraft.Empennage.HorizontalTail.WingSkin.Vne^2      WingCore      Mission1.Aircraft.Empennage.HorizontalTail.WingCore.W >= 2*(g*FoamHD.rho*Mission1.Aircraft.Empennage.HorizontalTail.WingCore.Abar*cave[:]^2*Mission1.Aircraft.Empennage.HorizontalTail.Planform.b/2*deta[:]).sum()     mh*(1 + 2/Mission1.Aircraft.Empennage.HorizontalTail.Planform.AR) <= 6.28     VerticalTail     Mission1.Aircraft.Empennage.VerticalTail.W/Mission1.Aircraft.Empennage.VerticalTail.mfac >= Mission1.Aircraft.Empennage.VerticalTail.WingSkin.W + Mission1.Aircraft.Empennage.VerticalTail.WingCore.W      Planform      Mission1.Aircraft.Empennage.VerticalTail.Planform.b^2 = Mission1.Aircraft.Empennage.VerticalTail.Planform.S*Mission1.Aircraft.Empennage.VerticalTail.Planform.AR      cave[:] = cbave[:]*Mission1.Aircraft.Empennage.VerticalTail.Planform.S/Mission1.Aircraft.Empennage.VerticalTail.Planform.b      Mission1.Aircraft.Empennage.VerticalTail.Planform.croot = Mission1.Aircraft.Empennage.VerticalTail.Planform.S/Mission1.Aircraft.Empennage.VerticalTail.Planform.b*cbar[0]      Mission1.Aircraft.Empennage.VerticalTail.Planform.cmac = Mission1.Aircraft.Empennage.VerticalTail.Planform.croot*Mission1.Aircraft.Empennage.VerticalTail.Planform.cbarmac     WingSkin      Mission1.Aircraft.Empennage.VerticalTail.WingSkin.W >= CFRPFabric.rho*Mission1.Aircraft.Empennage.VerticalTail.Planform.S*2*Mission1.Aircraft.Empennage.VerticalTail.WingSkin.t*g      Mission1.Aircraft.Empennage.VerticalTail.WingSkin.t >= tmin      CFRPFabric.tau >= 1/Mission1.Aircraft.Empennage.VerticalTail.WingSkin.Jtbar/Mission1.Aircraft.Empennage.VerticalTail.Planform.croot^2/Mission1.Aircraft.Empennage.VerticalTail.WingSkin.t*Mission1.Aircraft.Empennage.VerticalTail.WingSkin.Cmw*Mission1.Aircraft.Empennage.VerticalTail.Planform.S*Mission1.Aircraft.Empennage.VerticalTail.WingSkin.rhosl*Mission1.Aircraft.Empennage.VerticalTail.WingSkin.Vne^2      WingCore      Mission1.Aircraft.Empennage.VerticalTail.WingCore.W >= 2*(g*FoamHD.rho*Mission1.Aircraft.Empennage.VerticalTail.WingCore.Abar*cave[:]^2*Mission1.Aircraft.Empennage.VerticalTail.Planform.b/2*deta[:]).sum()     TailBoom     I[:] <= PI*t[:]*d[:]^3/8     Sy[:] <= 2*I[:]/d[:]     dm[:] >= PI*CFRPFabric.rho*d[:]*Mission1.Aircraft.Empennage.TailBoom.deta*t[:]*kfac*Mission1.Aircraft.Empennage.TailBoom.l     Mission1.Aircraft.Empennage.TailBoom.W/Mission1.Aircraft.Empennage.TailBoom.mfac >= g*dm[:].sum()     t[:] >= tmin     Mission1.Aircraft.Empennage.TailBoom.S = Mission1.Aircraft.Empennage.TailBoom.l*PI*d[0]     Mission1.Aircraft.Empennage.TailBoom.b = 2*Mission1.Aircraft.Empennage.TailBoom.l    Mission1.Aircraft.Empennage.W/Mission1.Aircraft.Empennage.mfac >= Mission1.Aircraft.Empennage.HorizontalTail.W + Mission1.Aircraft.Empennage.VerticalTail.W + Mission1.Aircraft.Empennage.TailBoom.W    Mission1.Aircraft.Empennage.TailBoom.l >= lh    Mission1.Aircraft.Empennage.TailBoom.l >= lv    Pylon    Mission1.Aircraft.Pylon.S >= 2*Mission1.Aircraft.Pylon.l*Mission1.Aircraft.Pylon.h   W_{zfw} >= Mission1.Aircraft.Fuselage.W + Mission1.Aircraft.Wing.W + Mission1.Aircraft.DF70.W + Mission1.Aircraft.Empennage.W + Mission1.Aircraft.Pylon.W + W_{pay} + W_{avn}   Vh <= Mission1.Aircraft.Empennage.HorizontalTail.Planform.S*lh/Mission1.Aircraft.Wing.Planform.S^2*Mission1.Aircraft.Wing.Planform.b   Vv = Mission1.Aircraft.Empennage.VerticalTail.Planform.S*lv/Mission1.Aircraft.Wing.Planform.S/Mission1.Aircraft.Wing.Planform.b   Mission1.Aircraft.Wing.Planform.CLmax/mw <= Mission1.Aircraft.Empennage.HorizontalTail.Planform.CLmax/mh   w_{antenna} <= Mission1.Aircraft.Empennage.VerticalTail.Planform.croot*Mission1.Aircraft.Empennage.VerticalTail.Planform.lam   Mission1.Aircraft.Empennage.VerticalTail.Planform.b >= l_{antenna}   Mission1.Aircraft.Empennage.TailBoom.l >= lh + Mission1.Aircraft.Empennage.HorizontalTail.Planform.croot   \mathcal{V}_{pay} <= (PI/1.5)*k_{nose}*R^3   \mathcal{V}_{body} >= \mathcal{V} + \mathcal{V}_{avn}   Mission1.Aircraft.DF70.h <= 2*R and AircraftLoading ===============  Cost ----  1  Constraints -----------   TailBoomBending    dx[:] = Mission1.Aircraft.Empennage.TailBoom.deta    qne[:]*Mission1.Aircraft.Empennage.HorizontalTail.Planform.S <= Mission1.AircraftLoading.TailBoomBending.F    \bar{EI}[:] <= CFRPFabric.E*I[:]/Mission1.AircraftLoading.TailBoomBending.F/l^2/2    Mr[:] >= \bar{M}[:-1]*Mission1.AircraftLoading.TailBoomBending.F*l    Mr[:]/Sy[:] <= CFRPFabric.sigma    Mission1.AircraftLoading.TailBoomBending.th = \theta[1]    Mission1.AircraftLoading.TailBoomBending.kappa >= \bar{\delta}[1]*Mission1.Aircraft.Empennage.HorizontalTail.Planform.CLmax*Mission1.AircraftLoading.TailBoomBending.Nsafety     Beam     \bar{M}[:-1] >= \bar{M}[1:] + 0.5*dx[:]*(\bar{S}[:-1] + \bar{S}[1:])     \bar{M}[1] >= Mission1.AircraftLoading.TailBoomBending.Beam.\bar{M}_{tip}     \theta[0] >= Mission1.AircraftLoading.TailBoomBending.Beam.\theta_{root}     \theta[1:] >= \theta[:-1] + 0.5*dx[:]*(\bar{M}[1:] + \bar{M}[:-1])/\bar{EI}[:]     \bar{\delta}[0] >= Mission1.AircraftLoading.TailBoomBending.Beam.\bar{\delta}_{root}     \bar{\delta}[1:] >= \bar{\delta}[:-1] + 0.5*dx[:]*(\theta[1:] + \theta[:-1])    TailBoomBending1    dx[:] = Mission1.Aircraft.Empennage.TailBoom.deta    qne[:]*Mission1.Aircraft.Empennage.VerticalTail.Planform.S <= Mission1.AircraftLoading.TailBoomBending1.F    \bar{EI}[:] <= CFRPFabric.E*I[:]/Mission1.AircraftLoading.TailBoomBending1.F/l^2/2    Mr[:] >= \bar{M}[:-1]*Mission1.AircraftLoading.TailBoomBending1.F*l    Mr[:]/Sy[:] <= CFRPFabric.sigma    Mission1.AircraftLoading.TailBoomBending1.th = \theta[1]    Mission1.AircraftLoading.TailBoomBending1.kappa >= \bar{\delta}[1]*Mission1.Aircraft.Empennage.VerticalTail.Planform.CLmax*Mission1.AircraftLoading.TailBoomBending1.Nsafety     Beam     \bar{M}[:-1] >= \bar{M}[1:] + 0.5*dx[:]*(\bar{S}[:-1] + \bar{S}[1:])     \bar{M}[1] >= Mission1.AircraftLoading.TailBoomBending1.Beam.\bar{M}_{tip}     \theta[0] >= Mission1.AircraftLoading.TailBoomBending1.Beam.\theta_{root}     \theta[1:] >= \theta[:-1] + 0.5*dx[:]*(\bar{M}[1:] + \bar{M}[:-1])/\bar{EI}[:]     \bar{\delta}[0] >= Mission1.AircraftLoading.TailBoomBending1.Beam.\bar{\delta}_{root}     \bar{\delta}[1:] >= \bar{\delta}[:-1] + 0.5*dx[:]*(\theta[1:] + \theta[:-1])    SparLoading    S[:-1] >= S[1:] + 0.5*deta[:]*b/2*(q[:-1] + q[1:])    M[:-1] >= M[1:] + 0.5*deta[:]*b/2*(S[:-1] + S[1:])    N = Mission1.AircraftLoading.SparLoading.Nsafety*Nmax    q[:] >= N*W/b*cbar[:]    S[11] >= Stip    M[11] >= Mtip    th[0] >= throot    th[1:] >= th[:-1] + 0.5*deta[:]*b/2*(M[1:] + M[:-1])/CFRPUD.E/I[:]    w[0] >= wroot    w[1:] >= w[:-1] + 0.5*deta[:]*b/2*(th[1:] + th[:-1])    M[:-1]/Sy[:] <= CFRPUD.sigma    Mission1.AircraftLoading.SparLoading.kappa >= w[11]/(b/2)    FuselageLoading    FuselageSkinL     M_h >= Mission1.AircraftLoading.FuselageLoading.FuselageSkinL.N_{max}*W_{cent}/4*l_{body}     Mission1.AircraftLoading.FuselageLoading.FuselageSkinL.\sigma_{Kevlar} >= M_h*R/Mission1.Aircraft.Fuselage.FuselageSkin.I     Mission1.AircraftLoading.FuselageLoading.FuselageSkinL.q >= W_{cent}*Mission1.AircraftLoading.FuselageLoading.FuselageSkinL.N_{max}/l_{body}     \kappa*l_{body}/2 >= Mission1.AircraftLoading.FuselageLoading.FuselageSkinL.q*(l_{body}/2)^4/8*Mission1.Aircraft.Fuselage.FuselageSkin.E*Mission1.Aircraft.Fuselage.FuselageSkin.I     FuselageLanding     Mission1.AircraftLoading.FuselageLoading.FuselageLanding.F >= W_{cent}*Mission1.AircraftLoading.FuselageLoading.FuselageLanding.N_{max}     a >= Mission1.AircraftLoading.FuselageLoading.FuselageLanding.F/m     \dot{\omega} >= a/(l_{body}/2)     M_G >= I_G*\dot{\omega}     Mission1.AircraftLoading.FuselageLoading.FuselageLanding.\sigma_{Kevlar} >= M_G*R/Mission1.Aircraft.Fuselage.FuselageSkin.I    TailBoomFlexibility    Fne >= 1 + mh*Mission1.AircraftLoading.TailBoomBending.th    sph1*mw*Fne/mh/Vh + deda <= 1    sph2 <= Vh*CLhmin/Mission1.Aircraft.Wing.Planform.CLmax    deda >= mw*Mission1.Aircraft.Wing.Planform.S/b/4/PI/lh    sph1 + sph2 >= SMcorr + CM/Mission1.Aircraft.Wing.Planform.CLmax.

Stacktrace

Traceback (most recent call last):
  File "c:\users\jenkins\workspace\ce_gpkit_pr_research_models\buildnode\windows10x64\optimizer\mosek\gpkit\constraints\set.py", line 86, in __init__
    self._update(subconstraint)
  File "c:\users\jenkins\workspace\ce_gpkit_pr_research_models\buildnode\windows10x64\optimizer\mosek\gpkit\constraints\set.py", line 114, in _update
    self.substitutions.update(constraint.substitutions)
  File "c:\users\jenkins\workspace\ce_gpkit_pr_research_models\buildnode\windows10x64\optimizer\mosek\gpkit\keydict.py", line 157, in update
    self[k] = v
  File "c:\users\jenkins\workspace\ce_gpkit_pr_research_models\buildnode\windows10x64\optimizer\mosek\gpkit\keydict.py", line 220, in __setitem__
    goodvals = ~np.isnan(value)
TypeError: ufunc 'isnan' not supported for the input types, and the inputs could not be safely coerced to any supported types according to the casting rule ''safe''

The above exception was the direct cause of the following exception:

Traceback (most recent call last):
  File "c:\users\jenkins\workspace\ce_gpkit_pr_research_models\buildnode\windows10x64\optimizer\mosek\gpkit\tests\helpers.py", line 59, in test
    testfn(name, import_dict, path)(self)
  File "c:\users\jenkins\workspace\ce_gpkit_pr_research_models\buildnode\windows10x64\optimizer\mosek\gpkit\tests\from_paths.py", line 48, in <lambda>
    lambda self: getattr(self, name)()))  # pylint:disable=undefined-variable
  File "c:\users\jenkins\workspace\ce_gpkit_pr_research_models\buildnode\windows10x64\optimizer\mosek\gpkit\tests\from_paths.py", line 37, in test_fn
    mod.test()
  File "C:\Users\jenkins\workspace\CE_gpkit_PR_research_models\buildnode\windows10x64\optimizer\mosek\jho\model\print_perf.py", line 194, in test
    M = Mission(DF70=True)
  File "c:\users\jenkins\workspace\ce_gpkit_pr_research_models\buildnode\windows10x64\optimizer\mosek\gpkit\constraints\model.py", line 71, in __init__
    CostedConstraintSet.__init__(self, cost, constraints, substitutions)
  File "c:\users\jenkins\workspace\ce_gpkit_pr_research_models\buildnode\windows10x64\optimizer\mosek\gpkit\constraints\costed.py", line 25, in __init__
    ConstraintSet.__init__(self, constraints, subs)
  File "c:\users\jenkins\workspace\ce_gpkit_pr_research_models\buildnode\windows10x64\optimizer\mosek\gpkit\constraints\set.py", line 88, in __init__
    raise badelement(self, i, constraint) from e
ValueError: Invalid ConstraintSet element '[<gpkit.Climb object containing 2 top-level constraint(s) and 933 variable(s)>, <gpkit.Cruise object containing 2 top-level constraint(s) and 364 variable(s)>, <gpkit.Loiter object containing 2 top-level constraint(s) and 612 variable(s)>, <gpkit.Cruise object containing 2 top-level constraint(s) and 364 variable(s)>]' <class 'list'> was between Aircraft
========

Cost
----
 1

Constraints
-----------
  Fuselage
   FuelTank
    Mission1.Aircraft.Fuselage.FuelTank.W >= f*W_{fuel-tot}
    \mathcal{V}/Mission1.Aircraft.Fuselage.FuelTank.m_{fac} >= W_{fuel-tot}/\rho_{fuel}

   FuselageSkin
    m >= S_{wet}*\rho_{kevlar}*Mission1.Aircraft.Fuselage.FuselageSkin.t
    Mission1.Aircraft.Fuselage.FuselageSkin.W >= m*Mission1.Aircraft.Fuselage.FuselageSkin.g
    Mission1.Aircraft.Fuselage.FuselageSkin.t >= t_{min}
    Mission1.Aircraft.Fuselage.FuselageSkin.I <= PI*R^3*Mission1.Aircraft.Fuselage.FuselageSkin.t
    I_G >= m*(4*R^2 + 4*R*Mission1.Aircraft.Fuselage.FuselageSkin.t + Mission1.Aircraft.Fuselage.FuselageSkin.t^2)
    l_{body} = l_{body}
    Mission1.Aircraft.Fuselage.FuselageSkin.E = Mission1.Aircraft.Fuselage.FuselageSkin.E
   k_{body} = l_{body}/R
   S_{wet} >= S_{body} + S_{nose} + S_{bulk}
   S_{body} >= 2PI*R*l_{body}
   S_{nose}^1.6 >= 2PI*R^2^1.6*(0.333 + 0.667*k_{nose}^1.6)
   S_{bulk} >= R^2*(0.0123*k_{bulk}^2 + 1.52*k_{bulk} + 0.502)
   \mathcal{V}_{body} <= PI*R^2*l_{body}
   Mission1.Aircraft.Fuselage.l <= 3*R*k_{body}*k_{nose}*k_{bulk}^0.3333333333333333
   Mission1.Aircraft.Fuselage.S >= PI*R^2
   \mathcal{V}_{body} >= \mathcal{V}
   Mission1.Aircraft.Fuselage.W/Mission1.Aircraft.Fuselage.m_{fac} >= Mission1.Aircraft.Fuselage.FuelTank.W + Mission1.Aircraft.Fuselage.FuselageSkin.W

  Wing
   Mission1.Aircraft.Wing.W/Mission1.Aircraft.Wing.mfac >= Mission1.Aircraft.Wing.WingSkin.W + Mission1.Aircraft.Wing.CapSpar.W + Mission1.Aircraft.Wing.WingCore.W

   Planform
    Mission1.Aircraft.Wing.Planform.b^2 = Mission1.Aircraft.Wing.Planform.S*Mission1.Aircraft.Wing.Planform.AR
    cave[:] = cbave[:]*Mission1.Aircraft.Wing.Planform.S/Mission1.Aircraft.Wing.Planform.b
    Mission1.Aircraft.Wing.Planform.croot = Mission1.Aircraft.Wing.Planform.S/Mission1.Aircraft.Wing.Planform.b*cbar[0]
    Mission1.Aircraft.Wing.Planform.cmac = Mission1.Aircraft.Wing.Planform.croot*Mission1.Aircraft.Wing.Planform.cbarmac
   WingSkin
    Mission1.Aircraft.Wing.WingSkin.W >= CFRPFabric.rho*Mission1.Aircraft.Wing.Planform.S*2*Mission1.Aircraft.Wing.WingSkin.t*g
    Mission1.Aircraft.Wing.WingSkin.t >= tmin
    CFRPFabric.tau >= 1/Mission1.Aircraft.Wing.WingSkin.Jtbar/Mission1.Aircraft.Wing.Planform.croot^2/Mission1.Aircraft.Wing.WingSkin.t*Mission1.Aircraft.Wing.WingSkin.Cmw*Mission1.Aircraft.Wing.Planform.S*Mission1.Aircraft.Wing.WingSkin.rhosl*Mission1.Aircraft.Wing.WingSkin.Vne^2

   CapSpar
    I[:]/Mission1.Aircraft.Wing.CapSpar.mfac <= 2*w[:]*t[:]*(hin[:]/2)^2
    dm[:] >= (CFRPUD.rho*2*w[:]*t[:] + 2*tshear[:]*CFRPFabric.rho*(hin[:] + 2*t[:]) + FoamHD.rho*w[:]*hin[:])*Mission1.Aircraft.Wing.Planform.b/2*deta[:]
    Mission1.Aircraft.Wing.CapSpar.W >= 2*dm[:].sum()*g
    w[:] <= wlim*cave[:]
    cave[:]*Mission1.Aircraft.Wing.Planform.tau >= hin[:] + 2*t[:]
    Sy[:]*(hin[:]/2 + t[:]) <= I[:]
    tshear[:] >= tmin

   WingCore
    Mission1.Aircraft.Wing.WingCore.W >= 2*(g*FoamHD.rho*Mission1.Aircraft.Wing.WingCore.Abar*cave[:]^2*Mission1.Aircraft.Wing.Planform.b/2*deta[:]).sum()
   mw*(1 + 2/Mission1.Aircraft.Wing.Planform.AR) >= 6.28

  DF70
   W_{DF70} <= Mission1.Aircraft.DF70.W/Mission1.Aircraft.DF70.m_{fac}
   P_{sl-max} = P_{sl-max}

  Empennage
   HorizontalTail
    Mission1.Aircraft.Empennage.HorizontalTail.W/Mission1.Aircraft.Empennage.HorizontalTail.mfac >= Mission1.Aircraft.Empennage.HorizontalTail.WingSkin.W + Mission1.Aircraft.Empennage.HorizontalTail.WingCore.W

    Planform
     Mission1.Aircraft.Empennage.HorizontalTail.Planform.b^2 = Mission1.Aircraft.Empennage.HorizontalTail.Planform.S*Mission1.Aircraft.Empennage.HorizontalTail.Planform.AR
     cave[:] = cbave[:]*Mission1.Aircraft.Empennage.HorizontalTail.Planform.S/Mission1.Aircraft.Empennage.HorizontalTail.Planform.b
     Mission1.Aircraft.Empennage.HorizontalTail.Planform.croot = Mission1.Aircraft.Empennage.HorizontalTail.Planform.S/Mission1.Aircraft.Empennage.HorizontalTail.Planform.b*cbar[0]
     Mission1.Aircraft.Empennage.HorizontalTail.Planform.cmac = Mission1.Aircraft.Empennage.HorizontalTail.Planform.croot*Mission1.Aircraft.Empennage.HorizontalTail.Planform.cbarmac
    WingSkin
     Mission1.Aircraft.Empennage.HorizontalTail.WingSkin.W >= CFRPFabric.rho*Mission1.Aircraft.Empennage.HorizontalTail.Planform.S*2*Mission1.Aircraft.Empennage.HorizontalTail.WingSkin.t*g
     Mission1.Aircraft.Empennage.HorizontalTail.WingSkin.t >= tmin
     CFRPFabric.tau >= 1/Mission1.Aircraft.Empennage.HorizontalTail.WingSkin.Jtbar/Mission1.Aircraft.Empennage.HorizontalTail.Planform.croot^2/Mission1.Aircraft.Empennage.HorizontalTail.WingSkin.t*Mission1.Aircraft.Empennage.HorizontalTail.WingSkin.Cmw*Mission1.Aircraft.Empennage.HorizontalTail.Planform.S*Mission1.Aircraft.Empennage.HorizontalTail.WingSkin.rhosl*Mission1.Aircraft.Empennage.HorizontalTail.WingSkin.Vne^2

    WingCore
     Mission1.Aircraft.Empennage.HorizontalTail.WingCore.W >= 2*(g*FoamHD.rho*Mission1.Aircraft.Empennage.HorizontalTail.WingCore.Abar*cave[:]^2*Mission1.Aircraft.Empennage.HorizontalTail.Planform.b/2*deta[:]).sum()
    mh*(1 + 2/Mission1.Aircraft.Empennage.HorizontalTail.Planform.AR) <= 6.28

   VerticalTail
    Mission1.Aircraft.Empennage.VerticalTail.W/Mission1.Aircraft.Empennage.VerticalTail.mfac >= Mission1.Aircraft.Empennage.VerticalTail.WingSkin.W + Mission1.Aircraft.Empennage.VerticalTail.WingCore.W

    Planform
     Mission1.Aircraft.Empennage.VerticalTail.Planform.b^2 = Mission1.Aircraft.Empennage.VerticalTail.Planform.S*Mission1.Aircraft.Empennage.VerticalTail.Planform.AR
     cave[:] = cbave[:]*Mission1.Aircraft.Empennage.VerticalTail.Planform.S/Mission1.Aircraft.Empennage.VerticalTail.Planform.b
     Mission1.Aircraft.Empennage.VerticalTail.Planform.croot = Mission1.Aircraft.Empennage.VerticalTail.Planform.S/Mission1.Aircraft.Empennage.VerticalTail.Planform.b*cbar[0]
     Mission1.Aircraft.Empennage.VerticalTail.Planform.cmac = Mission1.Aircraft.Empennage.VerticalTail.Planform.croot*Mission1.Aircraft.Empennage.VerticalTail.Planform.cbarmac
    WingSkin
     Mission1.Aircraft.Empennage.VerticalTail.WingSkin.W >= CFRPFabric.rho*Mission1.Aircraft.Empennage.VerticalTail.Planform.S*2*Mission1.Aircraft.Empennage.VerticalTail.WingSkin.t*g
     Mission1.Aircraft.Empennage.VerticalTail.WingSkin.t >= tmin
     CFRPFabric.tau >= 1/Mission1.Aircraft.Empennage.VerticalTail.WingSkin.Jtbar/Mission1.Aircraft.Empennage.VerticalTail.Planform.croot^2/Mission1.Aircraft.Empennage.VerticalTail.WingSkin.t*Mission1.Aircraft.Empennage.VerticalTail.WingSkin.Cmw*Mission1.Aircraft.Empennage.VerticalTail.Planform.S*Mission1.Aircraft.Empennage.VerticalTail.WingSkin.rhosl*Mission1.Aircraft.Empennage.VerticalTail.WingSkin.Vne^2

    WingCore
     Mission1.Aircraft.Empennage.VerticalTail.WingCore.W >= 2*(g*FoamHD.rho*Mission1.Aircraft.Empennage.VerticalTail.WingCore.Abar*cave[:]^2*Mission1.Aircraft.Empennage.VerticalTail.Planform.b/2*deta[:]).sum()

   TailBoom
    I[:] <= PI*t[:]*d[:]^3/8
    Sy[:] <= 2*I[:]/d[:]
    dm[:] >= PI*CFRPFabric.rho*d[:]*Mission1.Aircraft.Empennage.TailBoom.deta*t[:]*kfac*Mission1.Aircraft.Empennage.TailBoom.l
    Mission1.Aircraft.Empennage.TailBoom.W/Mission1.Aircraft.Empennage.TailBoom.mfac >= g*dm[:].sum()
    t[:] >= tmin
    Mission1.Aircraft.Empennage.TailBoom.S = Mission1.Aircraft.Empennage.TailBoom.l*PI*d[0]
    Mission1.Aircraft.Empennage.TailBoom.b = 2*Mission1.Aircraft.Empennage.TailBoom.l
   Mission1.Aircraft.Empennage.W/Mission1.Aircraft.Empennage.mfac >= Mission1.Aircraft.Empennage.HorizontalTail.W + Mission1.Aircraft.Empennage.VerticalTail.W + Mission1.Aircraft.Empennage.TailBoom.W
   Mission1.Aircraft.Empennage.TailBoom.l >= lh
   Mission1.Aircraft.Empennage.TailBoom.l >= lv

  Pylon
   Mission1.Aircraft.Pylon.S >= 2*Mission1.Aircraft.Pylon.l*Mission1.Aircraft.Pylon.h
  W_{zfw} >= Mission1.Aircraft.Fuselage.W + Mission1.Aircraft.Wing.W + Mission1.Aircraft.DF70.W + Mission1.Aircraft.Empennage.W + Mission1.Aircraft.Pylon.W + W_{pay} + W_{avn}
  Vh <= Mission1.Aircraft.Empennage.HorizontalTail.Planform.S*lh/Mission1.Aircraft.Wing.Planform.S^2*Mission1.Aircraft.Wing.Planform.b
  Vv = Mission1.Aircraft.Empennage.VerticalTail.Planform.S*lv/Mission1.Aircraft.Wing.Planform.S/Mission1.Aircraft.Wing.Planform.b
  Mission1.Aircraft.Wing.Planform.CLmax/mw <= Mission1.Aircraft.Empennage.HorizontalTail.Planform.CLmax/mh
  w_{antenna} <= Mission1.Aircraft.Empennage.VerticalTail.Planform.croot*Mission1.Aircraft.Empennage.VerticalTail.Planform.lam
  Mission1.Aircraft.Empennage.VerticalTail.Planform.b >= l_{antenna}
  Mission1.Aircraft.Empennage.TailBoom.l >= lh + Mission1.Aircraft.Empennage.HorizontalTail.Planform.croot
  \mathcal{V}_{pay} <= (PI/1.5)*k_{nose}*R^3
  \mathcal{V}_{body} >= \mathcal{V} + \mathcal{V}_{avn}
  Mission1.Aircraft.DF70.h <= 2*R and AircraftLoading
===============

Cost
----
 1

Constraints
-----------
  TailBoomBending
   dx[:] = Mission1.Aircraft.Empennage.TailBoom.deta
   qne[:]*Mission1.Aircraft.Empennage.HorizontalTail.Planform.S <= Mission1.AircraftLoading.TailBoomBending.F
   \bar{EI}[:] <= CFRPFabric.E*I[:]/Mission1.AircraftLoading.TailBoomBending.F/l^2/2
   Mr[:] >= \bar{M}[:-1]*Mission1.AircraftLoading.TailBoomBending.F*l
   Mr[:]/Sy[:] <= CFRPFabric.sigma
   Mission1.AircraftLoading.TailBoomBending.th = \theta[1]
   Mission1.AircraftLoading.TailBoomBending.kappa >= \bar{\delta}[1]*Mission1.Aircraft.Empennage.HorizontalTail.Planform.CLmax*Mission1.AircraftLoading.TailBoomBending.Nsafety

   Beam
    \bar{M}[:-1] >= \bar{M}[1:] + 0.5*dx[:]*(\bar{S}[:-1] + \bar{S}[1:])
    \bar{M}[1] >= Mission1.AircraftLoading.TailBoomBending.Beam.\bar{M}_{tip}
    \theta[0] >= Mission1.AircraftLoading.TailBoomBending.Beam.\theta_{root}
    \theta[1:] >= \theta[:-1] + 0.5*dx[:]*(\bar{M}[1:] + \bar{M}[:-1])/\bar{EI}[:]
    \bar{\delta}[0] >= Mission1.AircraftLoading.TailBoomBending.Beam.\bar{\delta}_{root}
    \bar{\delta}[1:] >= \bar{\delta}[:-1] + 0.5*dx[:]*(\theta[1:] + \theta[:-1])

  TailBoomBending1
   dx[:] = Mission1.Aircraft.Empennage.TailBoom.deta
   qne[:]*Mission1.Aircraft.Empennage.VerticalTail.Planform.S <= Mission1.AircraftLoading.TailBoomBending1.F
   \bar{EI}[:] <= CFRPFabric.E*I[:]/Mission1.AircraftLoading.TailBoomBending1.F/l^2/2
   Mr[:] >= \bar{M}[:-1]*Mission1.AircraftLoading.TailBoomBending1.F*l
   Mr[:]/Sy[:] <= CFRPFabric.sigma
   Mission1.AircraftLoading.TailBoomBending1.th = \theta[1]
   Mission1.AircraftLoading.TailBoomBending1.kappa >= \bar{\delta}[1]*Mission1.Aircraft.Empennage.VerticalTail.Planform.CLmax*Mission1.AircraftLoading.TailBoomBending1.Nsafety

   Beam
    \bar{M}[:-1] >= \bar{M}[1:] + 0.5*dx[:]*(\bar{S}[:-1] + \bar{S}[1:])
    \bar{M}[1] >= Mission1.AircraftLoading.TailBoomBending1.Beam.\bar{M}_{tip}
    \theta[0] >= Mission1.AircraftLoading.TailBoomBending1.Beam.\theta_{root}
    \theta[1:] >= \theta[:-1] + 0.5*dx[:]*(\bar{M}[1:] + \bar{M}[:-1])/\bar{EI}[:]
    \bar{\delta}[0] >= Mission1.AircraftLoading.TailBoomBending1.Beam.\bar{\delta}_{root}
    \bar{\delta}[1:] >= \bar{\delta}[:-1] + 0.5*dx[:]*(\theta[1:] + \theta[:-1])

  SparLoading
   S[:-1] >= S[1:] + 0.5*deta[:]*b/2*(q[:-1] + q[1:])
   M[:-1] >= M[1:] + 0.5*deta[:]*b/2*(S[:-1] + S[1:])
   N = Mission1.AircraftLoading.SparLoading.Nsafety*Nmax
   q[:] >= N*W/b*cbar[:]
   S[11] >= Stip
   M[11] >= Mtip
   th[0] >= throot
   th[1:] >= th[:-1] + 0.5*deta[:]*b/2*(M[1:] + M[:-1])/CFRPUD.E/I[:]
   w[0] >= wroot
   w[1:] >= w[:-1] + 0.5*deta[:]*b/2*(th[1:] + th[:-1])
   M[:-1]/Sy[:] <= CFRPUD.sigma
   Mission1.AircraftLoading.SparLoading.kappa >= w[11]/(b/2)

  FuselageLoading
   FuselageSkinL
    M_h >= Mission1.AircraftLoading.FuselageLoading.FuselageSkinL.N_{max}*W_{cent}/4*l_{body}
    Mission1.AircraftLoading.FuselageLoading.FuselageSkinL.\sigma_{Kevlar} >= M_h*R/Mission1.Aircraft.Fuselage.FuselageSkin.I
    Mission1.AircraftLoading.FuselageLoading.FuselageSkinL.q >= W_{cent}*Mission1.AircraftLoading.FuselageLoading.FuselageSkinL.N_{max}/l_{body}
    \kappa*l_{body}/2 >= Mission1.AircraftLoading.FuselageLoading.FuselageSkinL.q*(l_{body}/2)^4/8*Mission1.Aircraft.Fuselage.FuselageSkin.E*Mission1.Aircraft.Fuselage.FuselageSkin.I

   FuselageLanding
    Mission1.AircraftLoading.FuselageLoading.FuselageLanding.F >= W_{cent}*Mission1.AircraftLoading.FuselageLoading.FuselageLanding.N_{max}
    a >= Mission1.AircraftLoading.FuselageLoading.FuselageLanding.F/m
    \dot{\omega} >= a/(l_{body}/2)
    M_G >= I_G*\dot{\omega}
    Mission1.AircraftLoading.FuselageLoading.FuselageLanding.\sigma_{Kevlar} >= M_G*R/Mission1.Aircraft.Fuselage.FuselageSkin.I

  TailBoomFlexibility
   Fne >= 1 + mh*Mission1.AircraftLoading.TailBoomBending.th
   sph1*mw*Fne/mh/Vh + deda <= 1
   sph2 <= Vh*CLhmin/Mission1.Aircraft.Wing.Planform.CLmax
   deda >= mw*Mission1.Aircraft.Wing.Planform.S/b/4/PI/lh
   sph1 + sph2 >= SMcorr + CM/Mission1.Aircraft.Wing.Planform.CLmax.

Standard Output

Standard Error