Multipoint Optmization

[Ouzboy]

Hello,
Im Ousmane SY, PhD student at ISAE-SUPAERO.

I'm using OpenAeroStruct to perform trade-off studies of High Aspect Ratio wing concept by considering differents aspects of the ASO problem. I'm facing somes issues regarding the convergence of the multipoint case (more than 1 cruise flights points).Another issue Im facing is the implementation of the climb phase.
I noticed that both aspect were tackled by Eytan J. Adler in their work named "Efficient Aerostructural Wing Optimization Considering
Mission Analysis" attached below".
Would it be possible for you to share with me the implementation and ASO script used for that work?
Thanks in advance
Adler2022a.pdf

[eytanadler]

Hi Ousmane SY,
Unfortunately I don't have any scripts in a format that are ready to share and it's hard to help debug without any information about your code or the convergence failure modes you're observing. Before looking at the climb phase, I'd suggest getting multipoint working. There is at least one page in the docs that should help.

For the climb phase, I can offer the OpenMDAO component in which I modified the Breguet range equation to include a flight path angle input:

import numpy as np
import openmdao.api as om

class ModifiedBreguetRange(om.ExplicitComponent):
    """
    Computes the fuel burn using the Breguet range equation modified to include flight path angle.

    Parameters
    ----------
    CL : float
        Total coefficient of lift
    CD : float
        Total coefficient of drag
    CT : float
        Specific fuel consumption for the entire aircraft (1/s)
    V : float
        Flight speed
    R : float
        The total range over which to estimate fuel burn (m)
    W0 : float
        The initial aircraft weight, excluding the wing structural weight (kg)
    W_wing : float
        Weight of the wing's structure (kg)
    gamma : float
        Flight path angle (deg)

    Returns
    -------
    fuelburn : float
        Computed fuel burn in kg based on the Breguet range equation.
    """
    def setup(self):
        self.add_input("CT", val=0.25, units="1/s")
        self.add_input("CL", val=0.7)
        self.add_input("CD", val=0.02)
        self.add_input("R", val=3000.0, units="m")
        self.add_input("V", val=200, units="m/s")
        self.add_input("W0", val=200.0, units="kg")
        self.add_input("W_wing", val=0., units="kg")
        self.add_input("gamma", val=0., units="deg")

        self.add_output("fuelburn", val=1.0, units="kg")

        self.declare_partials("*", "*", method="cs")
        self.set_check_partial_options(wrt="*", method="cs", step=1e-30)

    def compute(self, inputs, outputs):

        CT = inputs["CT"]
        R = inputs["R"]
        V = inputs["V"]
        W0 = inputs["W0"]
        Ws = inputs["W_wing"]
        gamma = inputs["gamma"] * np.pi / 180

        CL = inputs["CL"]
        CD = inputs["CD"]

        equiv_L_D = 1 / (np.cos(gamma) * CD / CL + np.sin(gamma))

        fuelburn = (W0 + Ws) * (1 - np.exp(-R * CT / V / equiv_L_D))
        fuelburn[fuelburn < 0] *= 0  # can't have negative fuel burn
        outputs["fuelburn"] = fuelburn