Re: Calculating cd using openvsp python api

[Ганс]

From my experiance, the best way it's to take CD0 (profile drag) from the Parasite Drag section, the CDi and the CL from VSPAero results. Use them for an Overall CD and the CL/CD ratio. If you faced the problem when the GUI has a bug where it's not possible to see the obtained results. You should open "xxxxxxx.polar" file using Notepad++ (recomended) from the folder where the model is saved.

понеділок, 21 липня 2025 р. о 13:16:42 UTC+3 Ishaan Khullar пише:

Hi, 

Thanks so much for your help, I have almost got it working. I just had a couple of questions: I wanted to calculate the L/D ratio and hence wanted the total drag coefficient rather than the induced drag coefficient is there any way of getting this coefficient? (The only way i can think of currently is using aerodynamic formulas (CD = CD0 + KCL^2)). Also I don't think it is generating the vsp file properly such that I can view it properly in the GUI. Here is my revised code based on your suggestions. Any more tips? Thanks so much in advance

import sys
import csv
import pandas as pd

# Add OpenVSP Python module path
sys.path.append(r"C:\Users\khull\OneDrive\Documents\OpenVsp\OpenVSP-3.43.0-win64\python\openvsp")
import vsp

vsp.SetVSPAEROPath(r"C:\Users\khull\OneDrive\Documents\OpenVsp\OpenVSP-3.43.0-win64\python\openvsp\openvsp")

# Analysis conditions
aoa = 3.0
mach = 0.1
altitude = 0.0

# Design parameter ranges

camber_list = [0, 0.02, 0.04, 0.06]

cloc_list = [0.3, 0.4, 0.5]
tc_list = [0.09, 0.12, 0.15]

# Wing constants
span = 11.0
root_chord = 1.625
tip_chord = 0.325
sref = 16.32
cref = 1.38
bref = span

df = pd.DataFrame(columns=["Camber", "Camber_Loc", "Thickness", "CL", "CD", "L/D"])

# Loop through param combinations
for m in camber_list:
    for p in cloc_list:
        for t in tc_list:
            print(f"\n Running: camber={m}%, cloc={p}, thickness={t}")

            try:
                vsp.ClearVSPModel()

                wing_id = vsp.AddGeom("WING")

                vsp.SetGeomName(wing_id, "ParametricWing")

                # Set wing shape
                print("no")

                vsp.SetParmVal(wing_id, "Sym_Planar_Flag", "Sym", vsp.SYM_XZ)
                print("HI1")
             

                vsp.SetParmVal(wing_id, "SectTess_U", "XSec_1", 20)
                vsp.SetParmVal(wing_id, "Span", "XSec_1", span)
                print("HI2")
                vsp.SetParmVal(wing_id, "Sweep", "XSec_1", 0.0)
                print("HI3")
                vsp.SetParmVal(wing_id, "Dihedral", "XSec_1", 1.5)
                print("HI4")
                vsp.SetParmVal(wing_id, "Twist", "XSec_1", -2.0)
                print("HI5")
                vsp.SetParmVal(wing_id, "Root_Chord", "XSec_1", root_chord)
                print("HI6")
                vsp.SetParmVal(wing_id, "Tip_Chord", "XSec_1", tip_chord)

                # Set airfoil type and parameters
                xsecsurf_id = vsp.GetXSecSurf(wing_id, 0)

                # Change root and tip shapes
                vsp.ChangeXSecShape(xsecsurf_id, 0, vsp.XS_FOUR_SERIES)
                vsp.ChangeXSecShape(xsecsurf_id, 1, vsp.XS_FOUR_SERIES)
                vsp.Update()

                # Get section IDs
                xsec_id_root = vsp.GetXSec(xsecsurf_id, 0)
                xsec_id_tip = vsp.GetXSec(xsecsurf_id, 1)

                # Set NACA parameters at root
                camber = vsp.GetXSecParm(xsec_id_root, "Camber")
                camber_loc = vsp.GetXSecParm(xsec_id_root, "CamberLoc")

                thickness = vsp.GetXSecParm(xsec_id_root, "ThickChord")

                vsp.SetParmVal(camber, m)
                vsp.SetParmVal(camber_loc, p)
                vsp.SetParmVal(thickness, t)

                # Set NACA parameters at tip
                camber = vsp.GetXSecParm(xsec_id_tip, "Camber")
                camber_loc = vsp.GetXSecParm(xsec_id_tip, "CamberLoc")

                thickness = vsp.GetXSecParm(xsec_id_tip, "ThickChord")

                vsp.SetParmVal(camber, m)
                vsp.SetParmVal(camber_loc, p)
                vsp.SetParmVal(thickness, t)
                vsp.Update()

                print("HI8")

                # Save geometry for checking
                vsp_filename = f"vspaero_geom_c{m}_cloc{p}_t{t}.vsp3"

                vsp.WriteVSPFile(vsp_filename, vsp.SET_ALL)

                    print(" Available results:")

                    history_res = vsp.FindLatestResultsID("VSPAERO_History")
                    cl_vals = vsp.GetDoubleResults(history_res, "CL")

                    cd_vals = vsp.GetDoubleResults(history_res, "CDi")

                    if cl_vals and cd_vals:
                        CL = cl_vals[0]
                        CD = cd_vals[0]
                        L_D = CL / CD if CD != 0 else 0

                       

                        print(f" Done: CL={CL:.4f}, CD={CD:.4f}, L/D={L_D:.2f}")

                        new_row = {"Camber": m, "Camber_Loc": p, "Thickness": t, "CL": CL, "CD": CD, "L/D": L_D}
                        df = pd.concat([df, pd.DataFrame([new_row])], ignore_index=True)    

                   
                       
                    else:
                        raise ValueError("Missing CL/CD values in VSPAERO results")

                except Exception as e:
                    print(f" Error for camber={m}, cloc={p}, thickness={t}: {e}")

            except Exception as e:
                print(f" Error for camber={m}, cloc={p}, thickness={t}: {e}")

print(df)

[Ганс]

I forgot, you should also add the some increase coefficient for the total CD. It's because the VSPAero doesn't takes account for the viscosity/friction between materials (air to "your aircraft material"). I add usually CDf = +0.025. It's just for conservatism. But you can use other values.

понеділок, 21 липня 2025 р. о 16:11:49 UTC+3 Ганс пише:

[Ганс]

I just realised that you do in Python and my answers are not helpful/useful for you :0. I'm sorry for this. Maybe somebody takes it account. 

понеділок, 21 липня 2025 р. о 16:16:55 UTC+3 Ганс пише: